AIRCRAFT ACCIDENT INVESTIGATION REPORT · aircraft accident investigation report crash after...

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AA2017-4 AIRCRAFT ACCIDENT INVESTIGATION REPORT PRIVATELY OWNED J A 4 0 6 0 July 18, 2017

Transcript of AIRCRAFT ACCIDENT INVESTIGATION REPORT · aircraft accident investigation report crash after...

Page 1: AIRCRAFT ACCIDENT INVESTIGATION REPORT · aircraft accident investigation report crash after takeoff privately owned piper pa-46-350p, ja4060 chofu city, tokyo metropolitan, japan

AA2017-4

AIRCRAFT ACCIDENT

INVESTIGATION REPORT

PRIVATELY OWNED

J A 4 0 6 0

July 18, 2017

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The objective of the investigation conducted by the Japan Transport Safety Board in accordance

with the Act for Establishment of the Japan Transport Safety Board and with Annex 13 to the

Convention on International Civil Aviation is to determine the causes of an accident and damage

incidental to such an accident, thereby preventing future accidents and reducing damage. It is not the

purpose of the investigation to apportion blame or liability.

Kazuhiro Nakahashi

Chairman

Japan Transport Safety Board

Note:

This report is a translation of the Japanese original investigation report. The text in Japanese shall

prevail in the interpretation of the report.

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AIRCRAFT ACCIDENT INVESTIGATION REPORT

CRASH AFTER TAKEOFF

PRIVATELY OWNED

PIPER PA-46-350P, JA4060

CHOFU CITY, TOKYO METROPOLITAN, JAPAN

AT AROUND 10:58 JST, JULY 26, 2015

July 7, 2017

Adopted by the Japan Transport Safety Board

Chairman Kazuhiro Nakahashi

Member Toru Miyashita

Member Toshiyuki Ishikawa

Member Yuichi Marui

Member Keiji Tanaka

Member Miwa Nakanishi

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SYNOPSIS

<Summary of the Accident>

On Sunday, July 26, 2015, at around 10:58 Japan Standard Time (JST: UTC + 9

hrs: unless otherwise stated, all times are indicated in JST using the 24-hour clock), a

privately owned Piper PA-46-350P, registered JA4060, crashed into a private house at

Fujimi Town in Chofu City, right after its takeoff from Runway 17 of Chofu Airport.

There were five people on board, consisting of a captain and four passengers. The captain

and one passenger died and three passengers were seriously injured. In addition, one

resident died and two residents had minor injuries.

The aircraft was destroyed and a fire broke out. The house where the aircraft had

crashed into were consumed in a fire and neighboring houses sustained damage due to

the fire and other factors.

<Probable Causes>

It is highly probable that this accident occurred as the speed of the Aircraft

decreased during takeoff and climb, which led the Aircraft to stall and crashed into a

residential area near Chofu Airport.

It is highly probable that decreased speed was caused by the weight of the Aircraft

exceeding the maximum takeoff weight, takeoff at low speed, and continued excessive

nose-up attitude.

As for the fact that the Captain made the flight with the weight of the Aircraft

exceeding the maximum takeoff weight, it is not possible to determine whether or not

the Captain was aware of the weight of the Aircraft exceeded the maximum takeoff

weight prior to the flight of the accident because the Captain is dead. However, it is

somewhat likely that the Captain had insufficient understanding of the risks of making

flights under such situation and safety awareness of observing relevant laws and

regulations.

It is somewhat likely that taking off at low speed occurred because the Captain

decided to take a procedure to take off at such a speed; or because the Captain reacted

and took off due to the approach of the Aircraft to the runway threshold.

It is somewhat likely that excessive nose-up attitude was continued in the state

that nose-up tended to occur because the position of the C.G. of the Aircraft was close to

the aft limit, or the Captain maintained the nose-up attitude as he prioritized climbing

over speed.

Adding to these factors, exceeding maximum takeoff weight, takeoff at low speed

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and continued excessive nose-up attitude, as the result of analysis using mathematical

models, it is somewhat likely that the decreased speed was caused by the decreased

engine power of the Aircraft; however, as there was no evidence of showing the engine

malfunction, it was not possible to determine this.

< Recommendations >

In this accident, small private aircraft crashed into a residential area and caused

injury to residents as well as damages to houses, however the Aircraft was flying with

exceeding the maximum takeoff weight and without satisfying the requirements for

performance prescribed in the flight manual, and over the past five years, there have

been two fatal accidents involving small private aircraft affected by inappropriate weight

and position of the center of gravity of the aircraft ( (1) Mooney M20C, JA3788, which

crashed when landing at Yao Airport in March 2016, and (2) Cessna 172N Ram, JA3814,

which veered off the runway of Otone Airfield, Kawachi Town, Inashiki-gun, Ibaraki

Prefecture, and made a fatal contact with a ground worker in August 2012). In view of

the result of these accident investigations, as operation safety of small private aircraft

needs to be improved, the Japan Transport Safety Board recommends the Minister of

Land, Infrastructure Transport and Tourism pursuant to Article 26 of the Act for

Establishment of the Japan Transport Safety Board to take the following measures:

(1) Promote pilots of small private aircraft to understand the importance to confirm

that requirements for performance prescribed in the flight manual are satisfied, in

addition to the importance to comply with maximum takeoff weight and limit for the

position of the center of gravity, as confirmation before departure, at the occasions like

specific pilot competency assessments and aviation safety seminars.

Enforce instructions and trainings to pilots of small private aircraft to plan the

actions in advance including to follow the emergency procedure prescribed in the flight

manual and confirm these actions thorough self-briefing by a pilot himself at the time

of preparation before departure. along with compliance with the speed and procedure

prescribed in the flight manual, as for the actions to the situation of degraded flight

performance due to lack of acceleration or decrease in speed during takeoff.

(2) Study and compile the cases of effective measures connecting entrance taxiways

to runway thresholds in order to make maximum use of runway length and inform

aerodrome providers and administrators of these case studies as maximum use of

runway length at takeoff, will allow a pilot to have a margin to make a decision during

takeoff roll and contribute to improving safety.

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Abbreviations used in this report are as follow:

AEIS: Aeronautical En-route Information Service

AIP: Aeronautical Information Publication

ALT: altitude

AOA: Angle of Attack

ASI: Air Safety Investigator

ATSB: Australian Transport Safety Bureau

BLWR : Blower

CAA: Civil Aviation Authority

CAS : Calibrated Air Speed

CHT: Cylinder Head Temperature

COND : condition

DME: Distance Measuring Equipment

deg: degree

EAS: Equivalent Air Speed

EMERG: emergency

FAA : Federal Aviation Administration

FAR: Federal Aviation Regulations

fpm: feet per minutes

ft: feet

gal: gallon

GPH: Gallons Per Hour

GS: Ground Speed

HP: Horse Power

Hz: Hertz

IAS: Indicated Air Speed

ILS: Instrument Landing System

IMG: image

in: inch

JAX : Japan Aerospace Exploration Agency

JST: Japan Standard Time

kt: knot

KTS : knots

lb: pound

MAX: Maximum

NTSB: National Transportation Safety Board

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POH: Pilot’s Operating Handbook

PSI: Pounds per Square Inch

RNAV: Area Navigation

RPM: Revolutions/Rotations Per Minute

RWY: Runway

SMS: Safety Management System

STC: Supplemental Type Certificate

TAS : True Air Speed

TC: Type Certification

TCA : Terminal Control Area

TIT: Turbine Inlet Temperature

VFR : Visual Flight Rules

VHF: Very High Frequency

VOR: VHF Omni-directional radio Range

VRB: variable

Vs: Stall Speed

Vx: best angle of climb speed

Vy: best rate of climb speed

WDI: Wind Direction Indicator

Unit Conversion Table

1 ft : 0.3048 m

1 in: 25.40 mm

1 nm: 1,852 m

1 lb: 0.4526 kg

[ºC]: ([ºF] – 32) × 5/9

1 US gal: 3.785 ℓ

1kt : 1.852 km/h (0.514 m/s)

1 inHg: 3,386 Pa

1 HP: 0.746 Kw

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TABLE OF CONTENT

1. PROCESS AND PROGRESS OF THE ACCIDENT INVESTIGATION ............ 1

1.1 Summary of the Accident .............................................................................. 1

1.2 Outline of the Accident Investigation............................................................ 1

1.2.1 Investigation Organization.................................................................. 1

1.2.2 Representatives of the Relevant State ................................................ 1

1.2.3 Implementation of the Investigation ................................................... 1

1.2.4 Comments from the Parties Relevant to the Cause of the Accident ....

.............................................................................................................. 2

1.2.5 Comments from the Relevant State .................................................... 2

2. FACTUAL INFORMATION ................................................................................ 2

2.1 History of the Flight ...................................................................................... 2

2.1.1 History of the Flight based on Video Images and Photo Images

Taken in the Cabin of the Aircraft ....................................................................... 3

2.1.2 Statements of passengers and eyewitnesses ....................................... 4

2.2 Injuries to Persons ......................................................................................... 8

2.3 Damage to the Aircraft .................................................................................. 8

2.3.1 Extent of Damage ................................................................................ 8

2.3.2 Damage to the Aircraft Components ................................................... 8

2.4 Situation of Accident Site and Other Damage .............................................. 8

2.4.1 Situation of Accident Site .................................................................... 8

2.4.2 Other Damage .................................................................................... 11

2.5 Personnel Information ................................................................................. 14

2.6 Aircraft Information .................................................................................... 14

2.6.1 Aircraft ............................................................................................... 14

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2.6.2 Engine and Propeller ......................................................................... 15

2.6.3 Maintenance Records ........................................................................ 15

2.6.4 Weight and Balance ........................................................................... 16

2.6.5 Fuel and Lubricating Oil ................................................................... 19

2.7 Outline of Piper PA-46-350P ....................................................................... 19

2.7.1 Outline ............................................................................................... 19

2.7.2 Function of Each System ................................................................... 22

2.7.3 Characteristics of PA-46-350P .......................................................... 23

2.8 Confirmation of Weight and Position of the C.G. ........................................ 24

2.8.1 “Confirmation before Departure by Pilot in Command” Stipulated

in Civil Aeronautic Act ....................................................................................... 24

2.8.2 Importance of Confirmation and Influence on Flight Performance of

Weight and the Position of the C.G. .................................................................. 25

2.9 Meteorological Information ......................................................................... 27

2.9.1 General Weather Forecasts ............................................................... 27

2.9.2 Aeronautical Weather Observation at the Airport ........................... 28

2.9.3 Temperature on the Runway ............................................................. 28

2.10 Aerodrome Information ............................................................................ 29

2.11 Details of Damage .................................................................................... 32

2.12 Medical Information ................................................................................. 33

2.13 Information on Fire, and Fire-fighting and Rescue Operations .............. 33

2.13.1 Information of Fire and Fire-fighting Operation .............................. 33

2.13.2 Information on Rescue Operation ..................................................... 33

2.14 Descriptions in Flight Manual of the Aircraft ......................................... 33

2.14.1 SECTION 2 LIMITATIONS (Excerpts) ............................................ 33

2.14.2 SECTION 3 EMERGENCY PROCEDURES (Excerpts) .................. 37

2.14.3 SECTION 4 NORMAL PROCEDURES (Excerpts) .......................... 38

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2.14.4 SECTION 5 PERFORMANCE (Excerpts) ........................................ 46

2.14. 5 SECTION 6 WEIGHT AND BALANCE (Excerpts) .......................... 48

2.15 Descriptions from the Maintenance Manual and Other Documents for

PA-46-350P ............................................................................................................ 51

2.15.1 TIO-510-AE2A Operator’s Manual by the Engine Manufacturer .......... 51

2.15.2 Propeller Owner’s Manual (Manual No. 115N) by the Propeller

Manufacturer ..................................................................................................... 52

2.16 Tests and Verifications Information ........................................................ 52

2.16.1 Teardown Inspection of the Engine................................................... 53

2.16.2 Teardown Inspection of the Propeller ............................................... 53

2.16.3 Teardown Inspection of the Magnetos .............................................. 53

2.16.4 Teardown Inspection of the Air Conditioner ..................................... 53

2.16.5 Verification of the Flight Route based on Images ............................. 54

2.16.5.1 The obtained visual materials ........................................................... 54

2.16.5.2 Speed during the takeoff roll ............................................................. 57

2.16.5.3 The changes of speed and height after the takeoff ........................... 58

2.16.5.4 Estimated flight route ....................................................................... 60

2.16.5.5 Sound analysis during the before-takeoff check ............................... 61

2.16.5.6 Sound analysis during the flight ....................................................... 61

2.16.5.7 Readings of the instrument panel in the images taken in the cabin of

the Aircraft ......................................................................................................... 62

2.16.6 Analysis of the Past Flight ................................................................ 64

2.17 Information concerning the Operating and Maintenance Condition of the

Aircraft .................................................................................................................. 65

2.18 Additional Information ............................................................................. 66

2.18.1 Takeoff Ground Roll Distance and Takeoff Distance in Calculation ...

............................................................................................................ 66

2.18.2 Regulations concerning Takeoff Weight ........................................... 66

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2.18.3 Statements concerning the Captain .................................................. 69

2.18.4 Premature Nose-up at the Time of Takeoff....................................... 70

2.18.5 Flying on the “Backside” ................................................................... 71

2.18.6 Safety Measures upon Landing and Takeoff of Aircraft at the Airport

Facilities ............................................................................................................ 73

2.18.7 Status of the Maintenance Management of the Aircraft .................. 77

2.18.8 Status of Implementation of the Airworthiness Directive ................ 77

2.18.9 International Standards, Oversea Regulations and Reference Cases .

............................................................................................................ 77

2.18.10 Actions Taken by the Civil Aviation Bureau (Japan) after the

Accident ............................................................................................................ 83

2.18.11 Actions Taken by the Aerodrome Provider and Administrator (Tokyo

Metropolitan Government) after the Accident .................................................. 84

3. ANALYSIS.......................................................................................................... 86

3.1 Qualifications of Personnel .......................................................................... 86

3.2 Airworthiness Certificate of the Aircraft .................................................... 86

3.3 Takeoff Weight and Balance of the Aircraft ................................................ 86

3.3.1 Confirmation of Weight and the position of C.G. before Departure .....

............................................................................................................ 86

3.3.2 The Captain’s Recognition about Takeoff Weight ............................ 87

3.3.3 Effects of Weight and Balance of the Aircraft................................... 87

3.3.4 Deviation from Allowable Ranges for Weight and the Position of the

C.G.. ............................................................................................................ 87

3.4 Flight of the Aircraft at the Time of the Accident ....................................... 88

3.4.1 Configuration of the Aircraft and Weather Conditions at the Time of

the Accident ........................................................................................................ 89

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3.4.2 Comparison of Flight profiles of the Aircraft between at the Time of

the Accident and based on flight manual .......................................................... 89

3.4.3 Takeoff Roll and Takeoff (10:57:13-10:57:41) ................................... 90

3.4.4 Speed at the Time of Takeoff ............................................................. 91

3.4.5 State of Climbing just after Takeoff (10:57:42-10:57:54) .................. 91

3.4.6 Risks of Takeoff Slower than Lift-off Speed ...................................... 92

3.4.7 Situation following climb immediately after Takeoff (10:57:55-

10:58:00) ............................................................................................................ 93

3.4.8 Situation from Height Reduction to Crash (from 10:58:00) ............. 93

3.5 Analysis based on Mathematical Model ...................................................... 95

3.5.1 Conditions at the Time of the Accident ............................................. 95

3.5.2 Mathematical Model of the Aircraft .................................................. 96

3.5.3 Comparison at Takeoff Ground Rolls ................................................ 97

3.5.4 Simulation at Takeoff and Climb ...................................................... 99

3.5.5 Flight during Shallow Descent ........................................................ 101

3.6 Analysis on Engine Power of the Aircraft ................................................. 102

3.6.1 Decreased Engine Power due to Flying Operation ......................... 102

3.6.2 Influence of Outside Air Temperature ............................................ 103

3.6.3 Possibility of Decreased Engine Power Due to Other Factors ........ 103

3.6.4 Verification Result of Engine Power ............................................... 106

3.6.5 Relation between Decrease of Engine Power and RPM .................. 106

3.6.6 Constant Decrease of Value Indicated on TIT Indicator ................ 107

3.7 Preparation for Partial Power Loss of an Engine ..................................... 108

3.8 Response of the Civil Aviation Bureau of the Ministry of Land,

Infrastructure, Transport and Tourism .............................................................. 108

3.9 Response of the Provider and Administrator of Chofu Airport (Tokyo

Metropolitan Government) .................................................................................. 109

3.10 Improvement of Safety ........................................................................... 110

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3.10.1 Ensuring Safety of Small Private Aircraft ...................................... 110

3.10.2 Improvement of Safety at Airports ................................................. 111

3.10.3 Measures for Increasing Safety in Small Private Aircraft

Maintenance ..................................................................................................... 112

4. CONCLUSION ................................................................................................. 112

4.1 Findings ..................................................................................................... 112

4.2 Probable Causes ......................................................................................... 116

5. SAFETY ACTIONS .......................................................................................... 117

5.1 Safety Actions Taken after the Accident ................................................... 117

5.1.1 Safety Actions Taken by the Civil Aviation Bureau of the Ministry of

Land, Infrastructure, Transport and Tourism ................................................ 117

5.1.2 Safety Actions Taken by the Provider and administrator of Chofu

Airport (Tokyo Metropolitan Government)...................................................... 119

5.2 Safety Actions Required by the Civil Aviation Bureau of the Ministry of

Land, Infrastructure, Transport and Tourism .................................................... 120

6. RECOMMENDATIONS ................................................................................. 121

Figure 1 Three View Drawings of Piper PA-46-350P ....................................... 123

Figure 2 Seating Chart .................................................................................... 124

Attachment 1-1 Photos taken from inside of the Aircraft (Right Wing and Flaps)

(i) ..................................................................................................... 125

Attachment 1-2 Photos taken from inside of the Aircraft (Right Wing and Flaps)

(ii) .................................................................................................... 126

Attachment 2 Inside of the Aircraft (instrument panel) .................................... 127

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Attachment 3-1 Performance Chart (Takeoff Ground Roll Distance — 0° Flaps) ....

.................................................................................................... 128

Attachment 3-2 Performance Chart (Takeoff Distance — 0° Flaps) ................... 129

Attachment 3-3 Performance Chart (Short Field Takeoff — Takeoff Ground Roll

Distance) .................................................................................................... 130

Attachment 3-4 Performance Chart (Short Field Takeoff — Takeoff Distance) . 131

Attachment 4 Wind data Around the Time of the Accident ............................... 132

Attachment 5 Flight Test Record (Extract) ........................................................ 133

Attachment 6 Teardown Inspections of Engine, Propeller and Others .............. 134

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1. PROCESS AND PROGRESS OF THE ACCIDENT

INVESTIGATION

1.1 Summary of the Accident

On Sunday, July 26, 2015 at around 10:58 Japan Standard Time (JST: UTC + 9

hrs: unless otherwise stated, all times are indicated in JST using the 24-hour clock), a

privately owned Piper PA-64-350P, registered JA4060, crashed into a private house at

Fujimi Town in Chofu City, right after its takeoff from Runway 17 of Chofu Airport

There were five people on board, consisting of the captain and four passengers.

The captain and one passenger died and three passengers were seriously injured. In

addition, one resident died and two residents had minor injuries.

The aircraft was destroyed and a fire broke out. Furthermore, the house where

the Aircraft crashed into were consumed in a fire, and neighboring houses sustained

damage due to the fire and other factors.

1.2 Outline of the Accident Investigation

1.2.1 Investigation Organization

(1) On July 26, 2015, the Japan Transport Safety Board (JTSB) designated an

investigator-in-charge and two other investigators to investigate this accident. JTSB

designated six more investigators on September 26, 2016.

(2) An expert advisor was appointed for the investigation of the following technical

matters with respect to this accident.

For investigation into flight analysis

The Japan Aerospace Exploration Agency

Aeronautical Technology Directorate

Flight Research Unit Kouhei Funabiki

(Appointed on April 18, 2016)

1.2.2 Representatives of the Relevant State

An accredited representative of the United States of America, as the State of

Design and Manufacturer involved in this accident, participated in the investigation.

1.2.3 Implementation of the Investigation

July 26, 2015: On-Site investigation and aircraft examination

July 27, 2015: Aircraft examination, on-site investigation and interviews

July 28, 2015: Aircraft examination, document inspection and interviews

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July 29, 2015: Aircraft examination and interviews

August 3, 2015: Interviews and aircraft examination

January 12 to 13, 2016: Teardown inspections on the engine and propellers

(at the engine Manufacturer’s factory)

April 18, 2016 to January 5, 2017: Flight analysis

October 7, 2016: Flight test using the type of aircraft

December 13 to 14, 2016: Progress meeting with the relevant state concerning

the flight analysis and failure mode (held at the

Aircraft manufacturer’s factory)

1.2.4 Comments from the Parties Relevant to the Cause of the Accident

Comments were not invited from the parties relevant to the cause of the accident

due to the death of the captain.

1.2.5 Comments from the Relevant State

Comments on the draft report were invited from the relevant state.

2. FACTUAL INFORMATION

2.1 History of the Flight

On July 26, 2015, a privately owned Piper PA-46-350P (hereinafter referred to as

“the Aircraft”), registered JA4060, took off from Chofu Airport (hereinafter referred to

as “the Airport”), with the pilot sitting in the left pilot seat and four passengers being on

board the cabin (see Appended Figure 2 Seat Arrangement Diagram).

The flight plan of the Aircraft is outlined below:

Flight rules: Visual flight rules

Departure aerodrome: Chofu Airport

Estimated off-block time: 10:45

Cruising speed: 140 kt

Cruising altitude: VFR

Route: Yokosuka

Destination aerodrome: Oshima Airport

Total estimated elapsed time: 1 hour 00 minute

Fuel load expressed in endurance: 5 hours 00 minute

Persons on board: 5 persons

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Purpose of the flight: Other*1

The history of the flight up to the accident is summarized as below, according to

the multiple images that were taken around Chofu Airport (video images and pictures

taken in the cabin of the Aircraft), the statements of a passenger who sat in the left seat

of the rear row (hereinafter, referred to as “Passenger A”), a passenger who sat in the

right seat of the rear row (hereinafter, referred to as “Passenger B”), a passenger who

sat in the right seat of the middle row (hereinafter, referred to as “Passenger C”), Chofu

Flight Service personnel in charge of communication, and the eyewitnesses.

2.1.1 History of the Flight based on Video Images and Photo Images Taken in the

Cabin of the Aircraft

On the day of the accident, multiple images that were taken around the Airport

contained records of the flight situation of the Aircraft. Using these images, the

verified flight history is outlined below.

Around 10:50 The Aircraft had a preflight check at Spot 20N.

Around 10:54 Started taxiing for departure.

Approximately 10:57:12 Started a takeoff roll from near the threshold of Runway 17.

10:57:35 Passed through the point 400 m from the threshold of and at

the center of Runway 17 with a speed of approximately 59 kt.

Around this point, fluctuations on pitch were observed.

10:57:38 Nose gear showed movements of lift off the ground at a speed

of approximately 65 kt at the point approximately 500 m from

the threshold of Runway 17.

10:57:41 The Aircraft took off at the point approximately 630 m from

the threshold of Runway 17 with a speed of approximately 73

kt. The Aircraft continued to slowly veer to left from about the

time of takeoff.

10:57:52 Retracted the landing gear at a ground altitude (hereinafter

referred to as “height”) about 70 to 80 ft. The aircraft was nose-

up attitude and the climbing angle was approximately 4º.

10:57:55 Reached a height approximately 90 ft with a speed of

approximately 67 kt, then transferred from climbing to

descending slowly. However, the attitude of the Aircraft

remained a nose-up.

*1 “Other” means a flight other than air transport service and aerial work service and does not fall

under a flight for test, air transportation and official use.

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10:57:55 to 58:00 Along with repeating nose-up and pitch-down three times

during a slow descent, the speed decreased to approximately

62 kt with fluctuation of speed.

10:58:00 At a height approximately 84 ft, the Aircraft banked to left and

started to descend like sliding to lower-left. At this time, the

speed was approximately 62 kt and pitch was still nose-up

attitude.

10:58:07 Crashed into a private house in Chofu City.

Figure 2.1.1 Estimated Flight Route

2.1.2 Statements of passengers and eyewitnesses

(1) Passenger A

Passenger A planned the flight to Oshima and invited his acquaintances. A test-

run of the engine was conducted in the apron. The air-condition system was switched

on after the engine started. The takeoff was a standing takeoff, which the Aircraft

stopped once on the runway and started the roll after increasing engine rpm to full.

Passenger A did not feel any anomalies during the takeoff roll. At the time of the

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takeoff, a sequence of the turn to lift the gear was the nose gear first, and then the

main landing gear.

After the takeoff, Passenger A thought that the wind was strong as the plane

was swaying from side to side. He was thinking the Aircraft climbed so slow for a

powerful Malibu, but did not feel strange as the climb was also slow and gradual in

his past flights on a Cessna.

After that, he did not feel any danger even though the plane was unstable and

swaying from side to side because he thought the swaying was caused by strong wind.

The Captain did not seem to be panicking. He thought that the Captain was making a

left turn in order to pass through a checkpoint. He felt bad when he heard a sound like

a stall warning sound.

Passenger A did not remember the contact with private houses or the details of

the crash. He did not even prepare for anything because nobody told him any danger

and he himself did not feel anything dangerous. He did not know how long he had lost

consciousness, but when he became aware of the situation, he was upside down and

hanging by his seatbelt in the upside-down Aircraft. Somebody called out, “Fire”. It

was difficult to unfasten his seatbelt, but somehow he managed to unfasten it and

crawl out from the plane. It was very hot around him, and he could see roofing tiles

beside him. He could not move and was lying down there for a while. He saw other

passengers fall down.

(2) Passenger B

Passenger B was on board as Passenger A, who was his acquaintance, invited

him to hire a small airplane and enjoy flying. Passenger B felt that Passenger A often

invites his acquaintances for flights from the way he spoke.

Passenger C and the passenger who died (hereinafter referred to as “Passenger

D”) were friends of Passenger B from school days. After receiving the invitation by

Passenger A, three of them were tempted to go somewhere by hired small airplane.

Flying to Oshima was Passenger A’s suggestion and because their purpose was to fly

on a small airplane and they thought agreed without caring about the destination.

Passenger B thinks that he did tell his name and date of birth in advance, but

he did not remember that Passenger A or the Captain asked him about his weight or

other information.

In the Aircraft, Passenger B took the right seat of the rear row. The Captain sat

in the left pilot seat, Passenger C sat in the right seat of the middle row and Passenger

D in the left seat of the middle row. Passenger A sat in the left seat of the rear row.

Everybody took their seats without thought or instruction. All personnel on board

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fastened seatbelts. Nobody sat in the right pilot seat. Passenger B thought that the

Captain made a warm up engine and other operation from the time being on board to

the departure. After the takeoff or after the start of the engine, cold air flowed out of

the air-conditioning system.

After the takeoff, a change in attitude occurred right away. The Aircraft rolled

to the right or left right after the takeoff, and while Passenger B was wondering

whether it was OK, it rolled to the opposite side and crashed in just a second. There

was a terrible collision sound. He lost consciousness for a while due to the strong

impact because of the crash. He did not know for how long he had lost consciousness,

but he woke up because of strong pain in his abdomen.

When Passenger B returned to consciousness, he saw a fire. He did not know

when the fire broke out. Passenger A escaped to the outside of the Aircraft first.

Finding that Passenger C had been blown off with his chair, Passenger B escaped

taking Passenger C with him. He looked for Passenger D, but could not find him. The

place where they escaped was the second floor of a house.

A resident screamed “Fire!”

Passenger B stepped down to the ground holding onto a vehicle crashed beneath

the plane. Thinking of the risk of an imminent explosion, he evacuated to a spot about

30 m in front of a house. Some residents poured water over him as he lay down on the

ground. Passenger B did not hear any electronic “bii”, “puu”, or “pingpong” kind of

sounds.

(3) Passenger C

Prior to the flight, Passenger A replied to Passenger D, who asked about the

weight, that the weight would be OK because maximum passenger of the Aircraft was

six persons and there would be only five on board including the Captain. Up to the

accident flight, Passenger C does not have any memory of Passenger A or the Captain

asking him about his weight and other information.

In the Aircraft, Passenger C sat in the left seat of the middle row. Passenger D

sat in the right seat of the middle row and Passenger A and Passenger B sat in the

seats in the rear row. Everybody took their seats without thought or instruction.

The Captain was conducting before-takeoff check by calling out each item as he

checked it. The Aircraft stopped once after entering the runway and then started to

roll.

After the takeoff, he was surprised as the Aircraft descended suddenly.

Passenger C felt that the Aircraft lost balance and the Captain was trying to recover

the control while saying something. Passenger C has no memory of whether he heard

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the buzzer sound or not at this time.

Passenger C remembered the shock of the crash but had no recollection of the

following events. When he returned to consciousness, the Aircraft was crashed and he

thought he had to escape promptly. Passenger C smelled the aviation fuel.

(4) Chofu Flight Service Personnel in charge of communications and others

The Captain was conducting the radio communication of the Aircraft. There

were no specific anomalies during the takeoff roll, but the lift off seemed to be slightly

delayed. After the takeoff, it took long time to gain height and the Aircraft veered to

the left. After that, the personnel confirmed that a black smoke and a pillar of fire

were rising up.

(5) Eyewitness A (a pilot of another aircraft who was going to takeoff following the

Aircraft at the Airport)

Eyewitness A felt that the takeoff ground roll was long. After the takeoff, he

thought that the Aircraft would crash because it could not gain height and if it stayed

as it was. Then he saw that black smoke was rising and knew it crashed.

(6) Eyewitness B (a pilot who was preparing to depart from the apron of the Airport)

Eyewitness B saw there was no anomaly during the takeoff roll of the Aircraft.

It gained little height after the takeoff. The engine sound was normal. However, as the

Aircraft seemed panting while climbing, he noticed that something was wrong. It

seemed that the Aircraft tried to climb but failed, and looked like it was “floating”. He

thought that it might have been a situation like when flying at high altitude, engine

would not respond to the throttle operation.

(7) Eyewitness C (a neighboring resident of the accident site)

When Eyewitness C was in the living room, she heard the sound of an aircraft,

but it was not the sound she usually heard. Then, she heard a “bang” sound.

Eyewitness C heard the sound of an aircraft hitting a neighboring house, but

did not feel shaking. She clearly knew that an aircraft crashed, and went to the

window to look outside soon after she heard the sound.

Eyewitness C could not see the whole of the Aircraft, but could see a part of the

empennage of the Aircraft. She did not feel anything hitting her house. Immediately,

black smoke and an orange pillar of fire rose up from the direction of the Aircraft. She

did not see anyone escape from the Aircraft. There was no fuel raining. She rushed to

the main entrance and tried to open the door, but could not. She pushed the door with

the weight of her full body, and could finally open it and ran out to escape. The hot

blast of air was so severe that the planter of the morning glory had melted. On her

way to escape, she heard a crackling sound like fire and an explosion many times.

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(8) Eyewitness D (a neighboring resident of the accident site)

When Eyewitness D rushed to the site from the south side, neighbors were

pouring water over the passengers who escaped from the Aircraft. The fire was strong.

Those passengers suffered severe burns. Eyewitness D could hear a sound like a fire

cracker from the site. Fire engines arrived immediately. The Aircraft was upside down

and the wings were stuck out toward the street, but the empennage was standing

slightly diagonally.

The accident occurred at around 10:58 on July 26, 2015, near Fujimi Town in

Chofu City (N35º39’44”, E139º39’1”).

2.2 Injuries to Persons

The Captain, one passenger and one resident were killed. Three passengers were

seriously injured. Two residents suffered minor injuries.

2.3 Damage to the Aircraft

2.3.1 Extent of Damage

Destroyed.

2.3.2 Damage to the Aircraft Components

(1) Fuselage: Burned except a tail section

(2) Wings: Left wing fractured and burned, right wing damaged

(3) Engine: Burned

(4) Propeller: Bent and burned

(5) Landing Gear: Nose landing gear and main landing gear damaged

2.4 Situation of Accident Site and Other Damage

2.4.1 Situation of Accident Site

The accident site was in a residential area located about 770 m at 148 degree-

direction (south south-east) from the threshold of Runway 35 at Chofu Airport. The

private house with an antenna damaged by the Aircraft is referred to as “House A,”

the houses at south-east side from House A are referred to as “House B,” “House C”

and “House D” in order. In addition, the house adjacent to the north-east side of House

D is referred to as “House E.” (See Figure 2.4.1-1.)

The Aircraft crashed upside down with the nose to north and the empennage to

south in the site of House D, which faces a 4 m-wide public road at the south-west side.

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The fire that broke out from the Aircraft after the crash consumed and House

D completely.

House E was partially burned due to the fire caused from the Aircraft. Other

houses next to House D suffered damage due to the fire spread or radiant heat.

Excluding the empennage, the fuselage of the Aircraft was almost all burned up

without any original shapes. The right horizontal stabilizer was sticking out on the

public road. A part of the burned main wings was found on the public road in front of

House D and near the west-side wall of House D left unburned.

Figure 2.4.1-1 Area Map of Crash Site

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Photo 2.4.1-1 Accident Site and the Aircraft (part 1)

Photo 2.4.1-2 Accident Site and the Aircraft (part 2)

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2.4.2 Other Damage

House D facing a public road at the south-side was a two-story house, but the

south-side of the house was burned down almost completely, leaving only the west-

side wall. For the other parts of the house, the roofs were lost to fire and only charred

columns were standing. A vehicle and a motorcycle which had been parked in the

garden of House D were consumed by the fire, being under the Aircraft. Houses C and

E adjacent to House D, which was burned down completely, caught fire or suffered

damage due to radiant heat with the level of damage varying.

On the rooftop of House A, about 6.5 m high above the ground, a TV antenna

was collapsed together with the support column of about 3.8 m in length. The column

was bent by about 13º at the point of mounting the antenna branch line at a height of

about 2 m from the rooftop. A Yagi antenna for UHF installed at about 3 m high from

the rooftop was bent downward by about 75º, a part of the antenna was detached and

dropped onto the rooftop, and the antenna itself was off from the column and was hung

by single cable line from the roof edge.

The roofing materials of House B from the side facing the public road were

scattered onto the roof of nearby House C.

Photo 2.4.2-1 House A to House D

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Photo 2.4.2-2 Collapsed TV Antenna of House A

Photo 2.4.2-3 Damaged Roof of House B and Roofing Materials Scattered on House C

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There were the roofing materials from damaged House B scattered over the

rooftop of House C, but there is no trace of contact with the Aircraft. Besides, the

windows at the south east side of the first and second floors were broken and traces of

burning were confirmed on the south side exterior wall. A vehicle and a bicycle parked

in the garden were burned. House E

had a closet on the first floor and the

exterior wall and roof of the second

floor damaged by the fire.

In addition, it was confirmed

that the houses adjacent to House D

had entrance doors and other parts

burned in such as the fire, coatings for

outdoor air conditioner units,

ventilation fans and electrical wire

covers melted and window glasses

damaged and cracked.

Photo 2.4.2-4 Fire Damage of House C

Photo 2.4.2-5 Fire Damaged Vehicle

of House C

Photo 2.4.2-6 Fire Damage of House E

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2.5 Personnel Information

Captain Male, Age 36

Commercial pilot certificate (airplane) August 16, 2006

Type rating for Single engine (land) May 16, 2005

Multiple engine (land) December 22, 2005

Specific pilot competence March 31, 2014

Expiration date of piloting capable period March 31, 2016

Instrument flight certificate (airplane) May 10, 2006

Flight instructor certificate (airplane) February 1, 2013

Class 1 aviation medical certificate

Valid until February 23, 2016

Total flight time about 1,300 hours

Flight time in the last 30 days about 19 hours

Total flight time on the type of aircraft about 120 hours

Flight time in the last 30 days 0 hours 29 minutes

2.6 Aircraft Information

2.6.1 Aircraft

Type Piper PA-46-350P

Serial number No. 4622011

Date of manufacture February 14, 1989

Certificate of airworthiness No. To-27-058

Valid until May 1, 2016

Category of airworthiness Airplane Normal N

Total flight time 2,284 hours 50 minutes

Flight time since the last periodical check (100-hour check on April 17, 2015)

23 hours 51 minutes

(See Figure 1 Three Angle View of Piper PA-46-350P)

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Photo 2.6.1 The Aircraft (provided by the administrating office of the Airport)

2.6.2 Engine and Propeller

(1) Engine

Type: Lycoming TIO-540-AE2A

Serial Number: RL-9350-61A

Date of manufacture: March 22, 2003

Total flight time: 1,001 hours 32 minutes

(2) Propeller

Type: Hartzell HC-I2YR-1BF/F8074K

Serial Number: HA 6

Date of manufacture: June 6, 1988

Total flight time: 1,541 hours 01 minutes

2.6.3 Maintenance Records

During the time from April 8 to 17, 2015, the Aircraft received preparation work

to take an inspection for the Airworthiness Certificate (renewal) and renewed the

airworthiness certificate on May 1, 2015. There are no records of any periodic

maintenance work after that date.

The records of past flight tests accompanying periodic maintenance work had

items of Stall Speed, Stall Warning Speed, TIT and data of engine performance and so

on. (See Annex 5: The Records of Flight Test (Excerpt))

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(1) Airframe

As major maintenance work other than periodic maintenance work, repair work

was conducted from January 24 to June 16, 2005, for damage in the airframe due to

an accident which occurred on October 27, 2004. From June 23, 2008 to November 1,

2010, a 1000-hour check was implemented based on the maintenance manual.

(2) Engine

The engine of the Aircraft was replaced on June 18, 2004. Engine checks

relating to a propeller strike were conducted in conjunction with the damage of the

airframe due to the accident that occurred on October 27, 2004 was repaired. As

another major check, the 400-hour check of the engine was implemented along with

the 1,000-hour check of the airframe.

(3) Propeller

The propeller of the Aircraft was replaced on June 14, 2005, along with the

repair works for the damage of the airframe due to the accident that occurred on

October 27, 2004.

The maintenance records after the propeller replacement contain no statements

about any adjustment of the low-pitch stop setting.

2.6.4 Weight and Balance

Regarding the estimated takeoff weight of the Aircraft, the empty weight ( the latest

records by the airworthiness certificate examination ), weight of five passengers, cloths

and the like, belongings , onboard baggage (wheel stoppers, extra oil, mooring rope, rag,

stepladder, fire extinguisher, life vest, bags, flight logbook, Emergency Locator

Transmitter, first aid kit and the like), estimated amount of fuel, and amount of fuel

consumed for ground test-run and taxiing were estimated as confirmed:

(1) Empty weight: Approximately 1,358 kg (approximately 2,994 lb)

(2) Weight (Captain): Approximately 58.5 kg (approximately 129.0 lb)

(3) Weight (passengers): Approximately 280 kg (approximately 617.3 lb) in

total (based on the total of estimated weight of

four passengers)

(4) Cloths and the like of the passengers on board:

Total: approximately 7 kg (approximately 15.4 lb)

(5) Belongings of the passengers, onboard baggage:

Approximately 27 kg (approximately 59.5 lb)

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(6) Estimated amount of fuel: Approximately 286 kg

(approximately 105 gal = approximately 630.5 lb)

(7) Fuel consumed for ground test-run and taxiing

Approximately 8.2 kg

(approximately 3 gal = approximately 18 lb)

Total approximately 2,008 kg (approximately 4,427 lb)

According to the above, the weight of the Aircraft at the time of the accident was

estimated to be approximately 2,008 kg. Concerning the quantity of fuel, as the last

flight before the accident on July 22, 2015, the Captain made the flight for about 30

minutes and the Aircraft had the fuel filled to almost full right before this flight, the

weight was estimated based on these figures calculated from the fuel consumption in

the flight manual.

Since the designated maximum takeoff weight (hereinafter referred to as “the

maximum takeoff weight”) as limitations described in the flight manual of the Aircraft

is 1,950 kg (4,300 lb), it is highly probable that the takeoff weight of the Aircraft at the

time of the accident was exceeding the maximum takeoff weight by approximately 58

kg.

Regarding the position of the center of gravity (hereinafter referred to as “C.G.”)

of the Aircraft at the time of the accident, because it could not be determined for some

baggage whether it was loaded in the front or aft component, it is estimated to be

between + 146.0 and +146.5 in aft from the reference line. This position of the C.G. is

only for reference because it is highly probable that the weight of the Aircraft exceeded

the maximum takeoff weight, but it is highly probable that the position of the C.G was

close to the aft limit of the allowable range (C.G. range from + 143.3 to +247.2 in)

corresponding to the maximum takeoff weight (hereinafter the allowable range in this

report is that which corresponds to the maximum takeoff weight). (See Table 2.6.4 The

Weight and the Position of the C.G. of the Aircraft and Figure 2.6.4 The Weight and

the Position of the C.G. of the Aircraft (at the time of the Accident))

Furthermore, at the examination after the accident, no calculation sheet or

similar item was found to show that the Captain calculated the weight and the position

of the C.G. prior to the departure.

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Table 2.6.4 The Weight and the Position of the C.G. of the Aircraft (at the time of the

Accident)

Weight

(kg)

Weight

(Lb)

Arm Aft

Of Datum

(Inches)

Moment

(In.-Lb)

Basic Empty Weight 1,358 2,994 134.94 404,010

Pilot and Front Passenger 66.4 146.4 133.50 19,544

Passengers (Center Seats) 150.5 331.8 177.00 58,729

Passengers (Rear Seats) 134.5 296.5 218.75 64,859

Baggage (Forward)※ 14.7 32.4

88.60 2,871

6.8 15.0 1,329

Baggage (Aft) ※ 6.8 15.0

248.23 3,723

14.7 32.4 8,043

Fuel 286 630.5 150.31 94,770

Fuel Allowance for Taxi, & Runup 8.2 18 150.31 2,706

Total 2,008 4,427 146.00 645,800

146.50 648,578

※ Regarding baggage in forward and aft compartments, two sets of

estimated values are shown here.

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Figure 2.6.4 Weight and Balance of the Aircraft (at the time of the accident)

2.6.5 Fuel and Lubricating Oil

Fuel was aviation gasoline 100LL and lubricating oil was Phillips X/C MIL-L-

22851.

2.7 Outline of Piper PA-46-350P

2.7.1 Outline

Piper PA-46-350P (hereinafter, referred to as “PA-46-350P”) is a single

reciprocating engine aircraft manufactured by Piper. This engine is a horizontal

opposed-6 cylinder type single reciprocating engine with a turbocharger and produces

350 HP as a maximum power. The propeller is all-metal and has variable pitch control

with two 80-inch diameter blades. The Aircraft is all-metal, has retractable landing

gear, and is low winged. It has a pressurized cabin with seating for six occupants

(including the captain and co-pilot) and two baggage spaces with one in front and one

at the back of the cabin. As main control system, it has aileron, elevator and rudder.

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As characteristics of PA-46-350P, it has high cruising speed of 205 kt (380 km/h)

for a single reciprocating engine, the pressurized cabin, capabilities to climb up to

25,000 ft and others.

The PA-46-350P is in the PA-46 series of Piper Aircraft and was developed

based on the Piper PA-46-310 aircraft, with increased maximum power and other

improvements in flight performance.

(1) Basic Information

Maximum Takeoff Weight: 1,950 kg (4,300 lb)

Maximum Engine Speed: 2,500 rpm

Maximum Power : 350 HP

Stall Speed: Flap 36º 58 kt (with landing gear down)

Flap 0º 69 kt (with landing gear retracted)

Fuel Total Capacity: 122 gal (333 kg)

(2) Structure of Aircraft

The fuselage is an all-metal, semi-monocoque structure with three basic

fuselage sections: the forward baggage section, the pressurized cabin section, and the

empennage.

The seating arrangement is six seats in total, including two seats in the front

row (left seat for a captain, right for a co-pilot), two seats in the middle row (facing

backward) and two seats in the rear row (facing forward). Cabin access is through the

door, located on the left aft of the fuselage.

The wings are low wings with sealed integral fuel tanks utilizing structural

portions of the wings and hold 122 gal of fuel in total (including 2 gal of unusable fuel).

Pitot tubes and retractable landing gear are installed in the underside of the wings.

All-metal high-lift devices (flaps) are furnished on parts of the rear portions of the

wings. The flap operates through a push rod by an electric motor-actuator. The flap

has four positions of full-up (0º), 10º, 20º, and full down (36º), and a pilot selects every

flap position by flap control lever located in the instrument panel.

The all-metal ailerons are operated by a cable system connected to the control

wheel.

The empennage is all-metal; the vertical tail has a rudder and rudder trim, and

the horizontal tail has elevators and elevator trims. The rudder is controlled with the

rudder pedal at the foot of the pilot seat via torque tube. The elevators are controlled

by the input of control column via a cable and rod.

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(3) Engine

The engine is Lycoming TIO-540-AE2A, a turbocharged horizontally opposed 6

cylinder air-cooled engine, which produces 350HP. The maximum rated output is 350

HP/2,500 rpm and manifold pressure is 42.0 inHg. It uses 100 or 100 LL octane rated

aviation fuel.

A starter, two magnetos, a propeller governor, two alternators, two vacuum

pumps, an air conditioner compressor, and two turbochargers are equipped as

accessories.

A turbocharger is equipped, one each at the right and left sides of the engine.

Turbochargers extract energy from engine cylinder exhaust gases and use this energy

to compress engine induction air (maximum manifold pressure 42.0 inHg up to 20,600

ft). This allows the engine to maintain rated manifold pressure at high altitude.

Engine induction air is compressed by the turbochargers, and the air temperature

increases. The elevated air temperature is reduced by air intercoolers equipped on

each side of the engine. This helps to cool the inside of the engine and improves engine

power and efficiency.

The engine is equipped with a fuel injection system. An engine-driven fuel pump

supplies pressured fuel to the fuel injection regulator, which measures airflows and

meters the correct portion of fuel. The flow divider then directs the pressured fuel to

each of the individual cylinder injector nozzles. After combustion of the fuel in

cylinders, the exhaust gases are flowed to the exhaust manifold, then to the turbine of

the turbocharger which extracts the exhaust energy to operate the compressor.

Oil temperature and pressure information can be checked on the pilot’s

instrument panel.

(4) Propeller

The propeller is a Hartzell HC-I2YR-1BF/F8074, all-metal, and has variable

pitch control with two blades of 80 inches in rotation diameter. The propeller governor,

mounted on the front left of the engine, regulates the pressure of engine oil that flows

through the propeller shaft to change propeller pitch angle. The propeller governor

automatically changes the angle of the propeller in order to maintain the engine rpm

selected by a pilot. A propeller control lever is connected to the propeller governor with

cables to obtain the engine rpm selected by a pilot.

On this investigation, the propeller performance table was provided by the

propeller manufacturer. Using this propeller performance table, correlating values

such as an engine power, propeller pitch angle, and thrust were obtained by calculation.

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(5) Engine Control

The engine is controlled by throttle lever, propeller control lever and mixture

control lever, which are located on the control quadrant on the lower central

instrument panel.

The throttle lever is used to control engine power. The throttle lever

incorporates a gear-up warning horn switch, and the warning sound continues to

rumble as a warning to a pilot, if the throttle lever was at the low power position before

the landing gear is extended to be locked.

A propeller control lever adjusts the rpm of the engine. Setting the lever at the

foremost position leads to maximize rpm (minimum propeller pitch angle) and setting

it at the rearmost position (toward a pilot) minimizes rpm (maximum propeller pitch

angle). A propeller governor automatically changes the propeller pitch in order to

maintain the engine rpm selected by a pilot.

A mixture control lever adjusts the ratio of fuel and air. Moving the lever to the

front in full makes the mixing ratio richer, and moving to the rear (to a pilot) in full

suspends the supply of fuel and stops the engine.

2.7.2 Function of Each System

(1) Retractable landing gear

When the landing gear is down and locked, a gear light on the instrument panel

light up. While the gear is being retracted, the gear light is off and the landing gear

warning on an annunciator panel is lit. Completing to retract the gear, the landing

gear warning turns off.

(2) Annunciator Panel

In the center of the instrument panel, the annunciator panel, containing

warning lights, is installed.

Figure 2.7.2 (2) Annunciator Panel

(3) Magnetos

The soundness of the magnetos is confirmed in before-takeoff check. The

checking procedures on PA-46-350P are to set the prop speed lever at the foremost

position, advance the throttle lever and set to reach 2,000 rpm. At this time, with the

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condition that the low pitch stop is restricting the propeller pitch angle, the magnetos

are alternatively switched to confirm that the power is decreased as one of the

magnetos is turned off and that rpm could not be maintained and has dropped.

(4) Air-conditioning system

The Aircraft is equipped with an air-conditioning system utilizing a vapor

cycle.*2 It has a switch with three positions to change the operation: air-conditioner

/ off / blower. During the operation of the air-conditioner, cooled air comes out from six

small, semispherical holes called “eyeballs”. The switch position and the operating

status of the air-conditioner are as follows:

At the “air-conditioner” position: to send air into the cabin through producing

cooled air by simultaneous operation of the compressor equipped on the left side of the

engine and the blower (fan for ventilation)

At the “off” position: to stop the operation of the air-conditioner and the blower

At the “blower” position: to operate only the blower (fan for ventilation)

(5) Low pitch stop

The low pitch stop is a mechanism to physically limit the minimum propeller pitch

angle (most fine pitch) for a variable pitch propeller. At almost all phases during a

flight, a propeller governor controls the propeller pitch angles to keep constant rpm.

When the propeller pitch angle reaches the minimum angle set by the low pitch

stop, the propeller pitch angle is placed under the restricted condition. If the engine

power is reduced with the throttle lever, it becomes impossible to maintain rpm with

a further reduction of propeller pitch angle, thus rpm reduces.

(6) Outline of TIT indicator

A TIT indicator equipped on the Aircraft is outlined as below.

The TIT indicator is used to monitor the temperature of exhaust gas from

cylinders at the inlet to the turbocharger. In addition, while monitoring the inlet

temperature of the turbocharger, adjusting the mixture ratio of air and fuel with a

mixture control lever enables cruising at the most economical or the maximum power.

2.7.3 Characteristics of PA-46-350P

(1) Statement of a pilot with experience of flying the PA-46-350P

A pilot who has experience of flying the PA-46-350P described the

*2 “A vapor cycle” is a mechanism of an air conditioner wherein; refrigerant gas compressed with a

compressor is liquidized through cooling; the liquidized refrigerant gas is injected into an evaporator to

vaporize the refrigerant gas; in the process of vaporization, heat is taken away from the surrounding

and an evaporator is cooled; the cooled surrounding air is sent into the cabin via a blower to cool the

inside of the cabin.

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characteristics of the PA-46-350P and its operation at the Airport as below.

To take off from the Airport, he normally sets the flap at 10º. Even for a takeoff

with a 10º flap, nose-up at a speed lower than 78 kt is inconceivable. As the PA-46-

350P can only gain height at a very slow rate for climbing after takeoff, careful

consideration is required regarding takeoff weight.

In general, when flying a familiar aircraft, when a situation like the number of

persons on board, remaining fuel and others are known, it can be presumed that

whether the maximum takeoff weight is exceeded or close to being exceeding based on

experience without an accurate calculation.

(2) Statement of a maintenance engineer with experience of maintenance work on

the PA-46-350P

A maintenance engineer who has experience of maintenance work on the PA-

46-350P stated as follows:

In maintenance work in Japan, adjustment of the low pitch stop is not conducted.

It is common to adjust rpm by adjusting the link between the speed lever and the

propeller governor.

2.8 Confirmation of Weight and Position of the C.G.

2.8.1 “Confirmation before Departure by Pilot in Command” Stipulated in Civil

Aeronautic Act

Civil Aeronautics Act (Act No. 231 enacted on 1952) stipulates “Confirmation

before Departure by Pilot in Command” as follows (excerpts):

“Confirmation before Departure”

Article 73-2

The pilot in command shall not start an aircraft, unless he/she has

confirmed that the aircraft has no problems for flight and the necessary

preparation for air navigation has been completed, pursuant to the provision of

Ordinances of the Ministry of Land, Infrastructure, Transport and Tourism.

Besides, Ordinance for Enforcement of the Civil Aeronautics Act stipulates as

follows (excerpts):

“Confirmation before Departure”

Article 164-14

(1) Matters that must be confirmed by the pilot in command pursuant to

Article 73-2 of the Act are as listed below:

(i) Maintenance status of a subject aircraft and its equipment

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(ii) Take-off weight, landing weight, location of the center of gravity, and

weight distribution

(Omitted)

(2) A pilot in command shall, in the case of confirming the matters listed

under item (i) of the preceding paragraph, conduct the inspection of aircraft

logbook and other records on maintenance services, inspection of the exterior of

aircraft and ground trial run of engines, and other elemental inspection of

aircraft.

2.8.2 Importance of Confirmation and Influence on Flight Performance of Weight

and the Position of the C.G.

(1) “Airplane operation textbook” supervised by the Civil Aviation Bureau (Japan

Civil Aviation Promotion Foundation, March 31, 2009, p.56) has the following

descriptions:

“A training plane should have a flight manual and a flight manual has a weight

and balance calculation sheet. For an actual flight, a calculation must be made, using

these tables and it must be confirmed that all values are within a specified range for

weight and balance and complies with specific conditions of said flight. Calculation

procedure is to determine an empty weight, weights of passengers on board, weights

of loads (goods, fuel and lubrication) and other weights, then calculate the moment

indexes on each weight (kg × m or lb × inch). Judge total of these moments being

within the range surrounded by allowable weight and balance envelop or not to decide.

If flying with C.G. positioning outside of this range, it could result to cause serious

risks.”

(2) Regarding weight and position of the C.G., and stall, “Flight Dynamics 1:

Propeller Plane” published by Japan Aeronautical Engineer Association (the 2nd

version issued and revised on September 15, 2006) has the following descriptions:

(Excerpts)

Section 14 Weight and Position of the C.G.

Because allowable ranges for weight and position of the C.G. e are strictly

restricted from the viewpoint of airframe strength and maneuverability, and these are

indicated as operating limitations on airworthiness, it is essential to confirm that these

are within allowable ranges at all phases of flight condition at the flight planning stage.

14.4 Loading limitations

The limit can be exceeded depending on how the aircraft is loaded. In this case,

to limit the weight and to keep the C.G. within the allowable range, reduce the weight

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or relocate the load.

a. Weight limitations

It should be noted that a small aircraft fully occupied with passengers and

with a fully loaded fuel can often exceed the maximum take-off weight

limitations.

(Omitted)

c. Location of C.G. at the aft limit of the allowable range

An aircraft fully occupied with passengers can easily exceed not just the

weight limitations but the aft limit of the allowable range. Therefore, measures

must be taken to keep the C.G. within the allowable range such as by leaving

one of the rear seats vacant or measuring the weight of all people who are

boarding and assigning lighter passengers to the rear seats.

If the C.G. of an aircraft is close to the aft limit of the allowable range,

stability and controllability of the aircraft can still be maintained through

careful piloting. However, the fore portion of the aircraft becomes lighter than

desired, causing such strong tendencies as unstable ground roll, excessive

rotation rate during takeoff, reduced stability at low airspeeds, risk of stall and

spin, and greater difficulty in recovering from spin.

(Omitted)

Important precautions regarding the loading of a small aircraft are

summarized below.

(a) An aircraft with fully loaded fuel shall not fly with the seats fully

occupied or the cargo weight limit reached.

(b) An aircraft with the seats fully occupied shall have restriction of

fuel on board, resulting in shortening flight time and distance.

(c) The position of C.G. shall be within the allowable range throughout

the flight including the fuel consumption during the flight.

(d) Throughout the flight, the pilot shall stay aware of the position of

the C.G. and shall have a correct understanding of the difference of

control characteristics when the position of the C.G. is closer to the fore

or aft limit of the allowable range.

Section 15 Types of Stalls and Maximum Flight Movement

15.1 Types of Stalls

The following characteristics should be understood as common characteristics

of stall: a stall occurs in nature when the angle of attack for wings exceeds the stall

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angle, the aileron is the first to lose effectiveness among the three components of the

main control system and the rudder is the last to lose effectiveness; when the engine

power is higher, the stall speed become slower but changes in attitude and height at

the time of entering a stall become larger; when the GC position is aft, it is easier to

enter a stall and hard to recover.

(Omitted)

The stall speed is slower for a power-on stall at high power than for a power-off

stall, but it is easier to enter a stall with excessive nose-up in order to gain a climbing

angle during climbing at the stage of climbing right after a takeoff. If the engine power

is decreased sharply or the engine fails suddenly at this stage, there is the risk of

suddenly entering a “Complete Stall,” from which it is almost impossible to recover, or

a spin.

When the position of the C.G. is close to the aft limit, the generation of pitch-

down moment is small even though coming close to the stall speed and a delay in

maneuvering at an initial stall results in the risk of developing a flat spin which is

hard to recover from.

Flat spin is also called a “horizontal spin” and is a type of spin which results in a

rapid loss of height while rotating and maintaining the horizontal position of the

airframe. Because the spin causes a stall condition for a horizontal tail and vertical

tail at the same time and the elevator and rudder loses the effectiveness completely,

recovery by maneuvering is not possible. Special caution is required as it is easier to

cause a spin when the position of the C.G. is aft or the engine on one side of a multiple

engine aircraft has failed.

2.9 Meteorological Information

2.9.1 General Weather Forecasts

The general weather forecast released by Forecast Department of Japan

Meteorological Agency at 10:44 on July 26, 2015 (the accident day), was as follows:

A high pressure system centered in the sea south of Japan covers East Japan.

The Kanto-Koushin region has clear weather in general. On the 26th, a high pressure

system covers the area and it is clear in general, but there will be rain or

thunderstorms in the afternoon along mountains due to the effects of a high rise in

daytime temperature and moist air, depending on the area.

The Tokyo region is clear on the 26th and will be cloudy at night.

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2.9.2 Aeronautical Weather Observation at the Airport

(1) METAR observations

Aeronautical weather observation for aerodrome routine meteorological report

and aeronautical special meteorological report of the Airport were as follows:

10:00 Wind direction VRB, Wind velocity 1 kt, Visibility 15 km,

Cloud: Amount 1/8, Type: Cumulus, Cloud base: 3,000 ft,

Temperature 33ºC, Dew point 22ºC

Altimeter setting (QNH) 29.86 inHg

11:00 Wind direction VRB, Wind velocity 2 kt, Visibility 15 km,

Cloud: Amount 1/8, Type: Cumulus, Cloud base: 3,000 ft,

Temperature 34ºC, Dew point 22ºC,

Altimeter setting (QNH) 29.84 inHg

11:03 Wind direction VRB, Wind velocity 3 kt, Visibility 15 km,

Cloud: Amount 1/8, Type: Cumulus, Cloud base: 3,000 ft,

Temperature 34ºC, Dew point 21ºC,

Altimeter setting (QNH) 29.84 inHg

(2) Observations of wind direction and wind velocity at the time of takeoff

The meteorological information system of the Airport automatically records the

instantaneous wind velocity and instantaneous wind direction every 3 seconds. From

10:57:12, when the Aircraft started the takeoff roll from near the threshold of Runway

17, to 10:58:09, immediately after the accident, the recorded instantaneous wind

direction was 139º to 243º and the recorded instantaneous wind velocity was 0 to 1 kt.

(See Attachment 4 Wind Data at the Time Relating to the Accident)

2.9.3 Temperature on the Runway

The temperature on the runway at the time of accident was not observed.

At around 14:00, August 11, 2015, the observation value of temperature on the

runway of the airport was as follows:

At the centerline of the runway near A2 Taxiway (the estimated lift-off point of

the Aircraft): 38.1 ºC (approximately 1.5m above the ground)

In addition, the observation values according to the METAR were as follows:

14:00 Wind direction: 010º, Wind velocity: 5 kt,

Cloud amount: 1/8 – 2/8, Type: Cumulus, Cloud base: 3,000 ft,

Temperature: 34ºC

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2.10 Aerodrome Information

The aerodrome is located at approximately 22km to the west from Tokyo Station.

Most of the ground is within Chofu City and a part extends over Mitaka City and Fuchu

City. Athletic fields

and parks surround

the Airport, but the

area enclosing these

have much housing.

The official name is

Tokyo Metropolitan

Chofu Airport, and

Tokyo Metropolitan

government

establishes and

administers the

management of the

Airport.

The Airport is at an altitude 139 ft at reference point, 800 m in length and 30 m

in width, the runway magnetic direction is 170.20º/350.20º, and both ends of the runway

have overrun areas, which are 60 m in length and 30 m in width. Regarding the lightings

and markings related to the runway and the overrun areas, Visual Approach Slope

Indicator System and Runway End Identifier Lights are installed.

Figure 2.10-2 Site Plan of the Airport

(1) History of the Airport

March 1973: Return of the entire Airport area from the US Army

March 1979: Commencement of a regular air service between

Chofu and Niijima-island

Figure 2.10-1 Location of the Airport

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December 1984: Commencement of a regular air service between

Chofu and Oshima-island

July 1992: Transferring the management of the Airport from

Civil Aviation Bureau to Tokyo Metropolitan

Government

Commencement of a regular air service between

Chofu and Kozushima-island

December 25, 1998: Government’s permission to allow Tokyo

Metropolitan Government to establish an aerodrome

March 31, 2001: Started the service of an official airport (metropolitan

airport for commuters air service)

April 1, 2006: Commencement of aeronautical information service

by Tokyo Metropolitan Government

April 2, 2013: Started the service of a new passenger terminal

June 18, 2013: Introduction of IFR to a remote island air routes

April 2, 2014: Commencement of a regular air service between

Chofu and Miyakejima-island

(2) Operation of the aerodrome

The aerodrome provider and administrator (Tokyo Metropolitan Government)

set forth the Safety Codes for Tokyo Metropolitan Chofu Airport (Safety edition) (2009-

KoToCho-43) (hereinafter referred to as “Safety Codes”) based on Article 47-2 of the

Civil Aeronautics Act and the Tokyo Metropolitan Airport Regulations (1962

Regulations 53) in order to administrate and manage the facilities of the Airport,

properly and safely.

In order to secure the safety of the Airport and to preserve the living

environment of the area, the aerodrome provider and administrator (Tokyo

Metropolitan Government) set the operating procedures of the Airport, rules on entry

to the Airport restricted area, the action plan at the time of an emergency, and the

procedures concerning obstacle management around the Airport based on Safety

Codes along with providing restrictions on operation hours, number of departure and

arrival, and flight purpose in line with the Tokyo Metropolitan Airport Regulations

and the operational guidelines based thereon.

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(3) Situation surrounding the aerodrome

In order to ensure the safety of aircraft for departure and arrival, it is restricted

to install, plant, or leave any structures, plants or any other objects which protrude

above the approach surface, transitional surface or horizontal surface (hereinafter

referred to as “restricted surfaces”), and the aerodrome provider and administrator

(Tokyo Metropolitan Government) supervises to prohibit any objects protruding above

the restricted surfaces from being installed through checking upon granting building

certifications,

conducting

patrols,

inspection and

others as

needed.

There are

no objects taller

than the

restricted

surfaces of the

Airport.

Figure 2.10-3 Restricted Surfaces Applied to the Airport

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2.11 Details of Damage

(1) Fuselage

The pilot seat and the cabin were lost to

fire without the original shapes left. The aft part

of the fuselage had damage but it kept its

original shape there was little damage to the

empennage.

(2) Wings

The left wing was broken from the part

connected to the fuselage and was lost to fire

without the original shape left. The right wing

nearly kept its original shape, but the damage

by fire and the impact was severe.

Regarding the aileron, the left side was

burnt completely and the right side kept its

original shape but the damage by fire and the

impact was severe.

The left side flap was completely lost to

fire, the right side flap kept its original shape but the damage by fire and the impact

was severe. The positional relation between the front edge of the flap and the flap-

truck weight reduction hole of the Aircraft corresponded to be at 10º of the flap.

The flap actuator was positioned at 0º of the flap.

(3) Engine

The engine nearly kept its original shape, but most of the non-metal parts were

burnt.

Regarding turbochargers, the centrifugal compressor of the left turbocharger

was melted, but the right turbocharger kept its original shape.

The engine accessories kept their original shapes, but the damage by fire and

the impact was severe.

(4) Propeller

One blade was broken and burnt, the tip of the other one was bent backward

and burnt. Propeller governor remained mounted on the engine, but the tip was

damaged.

(5) Landing gears

The left landing gear was detached from the airframe but other landing gears

were found as being retracted. All landing gears were damaged and burnt.

Photo 2.11 Leading Edge of

Flap and Flap Truck

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2.12 Medical Information

According to the result of the legal autopsy carried out by the Tokyo Metropolitan

Police Department on June 26, 2015, the following were found:

(1) The captain tested negative for alcohol and drugs.

(2) Cause of death for the captain, Passenger D and the resident of House D were

death by fire.

2.13 Information on Fire, and Fire-fighting and Rescue Operations

Based on the report provided by Chofu Fire Department, the following were found:

2.13.1 Information of Fire and Fire-fighting Operation

On the day of the accident, at around 11:02, Chofu Fire Department received a

call for firefighters from a resident near the accident site, saying that he/she looked

outside because of a vehicle crashing sound and saw flames rising. Ambulances and

fire engines arrived at the site at 11:08 and started the fire-fighting operation at 11:09.

The fire was extinguished at 18:56 (total fire extinguishment confirmed). Because of

this fire, a total of 102 fire engines and ambulance cars were dispatched.

2.13.2 Information on Rescue Operation

Passengers A, B, and C, who escaped from the Aircraft, were evacuated to the

front of the house located about 30 m from the site. Neighboring residents were

pouring water over these three as a first aid. After the arrival of firefighters of Chofu

Fire Department on the site, they were transported to hospitals by ambulance. The

firefighters, while fighting the fire, rescued two survivors in the vicinity of the cabin

of the Aircraft around 11:24 to 27, and found one survivor near the courtyard on the

first floor of House D, and transported them by ambulance from the site.

The three persons carried out were the captain, Passenger D and the resident

of House D, all of whom were confirmed dead at hospital. Adding to these, two

residents of Houses C and D suffered injuries.

2.14 Descriptions in Flight Manual of the Aircraft

Descriptions in PILOT’S OPERATING HANDBOOK AND FAA APPROVED

AIRPLANE FLIGHT MANUAL of the Aircraft include the followings:

2.14.1 SECTION 2 LIMITATIONS (Excerpts)

2.7 POWER PLANT LIMITATIONS

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(d) Engine Operating Limits

(1) Maximum Engine Speed 2500 RPM

(2) Maximum Oil Temperature 245 °F

(3) Maximum Cylinder Head Temperature 500 °F

(4) Maximum Turbine Inlet Temperature 1750 °F

(5) Maximum Manifold Pressure

(inches of mercury)

To 20,600 feet 42

20,600 to 25,000 feet 42 - 1.6 per

1000 foot increase

(6) Minimum Manifold Pressure (IN. HG.)

Above 23,000 feet 23

(7) Minimum Propeller Speed (RPM)

Above 23,000 feet 2400

(j) Propeller Diameter (inches)

Minimum 79

Maximum 80

(k) Blade Angle Limits

Low Pitch Stop 17.6° +/- 0.2°

High Pitch Stop 40.5° +/- 0.5°

2.13 WEIGHT LIMITS

(a) Maximum Ramp Weight 4318 LB

(b) Maximum Takeoff Weight 4300 LB

(c) Maximum Landing Weight 4100 LB

(d) Maximum Zero Fuel Weight 4100 LB

NOTE

Refer to Section 5 (Performance) for maximum

weight as limited by performance.

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2.15 CENTER OF GRAVITY LIMITS

Weight Forward Limit Rearward Limit

Pounds Inches Aft of Datum Inches Aft of Datum

4300 143.3 147.1

4100 139.1 147.1

4000 137.0 146.5

2450 (and less) 130.7 137.6

2400 137.3

NOTE

Straight line variation between points given.

The datum used is 100.0 inches ahead of the

forward pressure bulkhead.

It is the responsibility of the airplane owner and

the pilot to ensure that the airplane is properly

loaded. See Section 6 (Weight and Balance) for

proper loading instructions.

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6.9 WEIGHT AND BALANCE DETERMINATION FOR FLIGHT

C.G. RANGE AND WEIGHT GRAPH

Figure 6-15

2.29 AIR CONDITIONING SYSTEM LIMITATIONS

AIR COND/BLWR switch in OFF or BLWR position for takeoffs and

landings.

NOTE

REC BLWR switch may be in HIGH or LOW

position.

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2.33 MAXIMUM SEATING CONFIGURATION

The maximum seating capacity is 6 (six) persons.

2.14.2 SECTION 3 EMERGENCY PROCEDURES (Excerpts)

3.3c. ENGINE POWER LOSS DURING TAKEOFF (3.9)

If sufficient runway remains for a normal landing, leave gear down and land

straight ahead.

If area ahead is rough, or if it is necessary to clear obstructions:

Landing Gear Selector ..............................................................................................UP

Mixture ................................................................................................ IDLE CUT-OFF

Emergency (EMERG) Fuel Pump ..........................................................................OFF

Fuel Selector ........................................................................................................... OFF

Battery Master (after gear retraction)................................................................... OFF

If sufficient altitude has been gained to attempt a restart:

Maintain Safe Airspeed.

Emergency (EMERG) Fuel Pump..................................................................Check ON

Fuel Selector .............................................................SWITCH to tank containing fuel

Mixture .......................................................................................................FULL RICH

Induction Air ........................................................................................... ALTERNATE

CAUTION

If power is not regained:

Prepare for power off landing.

3.9 ENGINE POWER LOSS DURING TAKEOFF (3.3c)

The proper action to be taken if loss of power occurs during takeoff will depend

If normal engine operation and fuel flow are not

reestablished, the emergency (EMERG) fuel

pump should be turned OFF. The lack of a fuel

flow indication could indicate a leak in the fuel

system. If fuel system leak is verified, switch

fuel selector to OFF.

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on the circumstances of the particular situation.

If sufficient runway remains to complete a normal landing, leave the landing

gear down and land straight ahead.

If the area ahead is rough, or if it is necessary to clear obstructions, move the

landing gear selector switch to the UP position and prepare for a gear up landing. If

time permits, move mixture control to idle cut-off, turn OFF the emergency

(EMERG) fuel pump, move the fuel selector to OFF and, after the landing gear is

retracted, turn battery master switch OFF.

If sufficient altitude has been gained to attempt a restart, maintain a safe

airspeed, turn the emergency (EMERG) fuel pump ON, and switch the fuel selector

to another tank containing fuel. Ensure the mixture is full RICH and move the

induction air lever to the ALTERNATE position.

If engine failure was caused by fuel exhaustion, power will not be regained

after switching fuel tanks until the empty fuel lines are filled. This may require up

to ten seconds.

If power is not regained, proceed with Power Off Landing procedure (refer to

paragraph 3.13).

2.14.3 SECTION 4 NORMAL PROCEDURES (Excerpts)

4.3 AIRSPEEDS FOR SAFE OPERATIONS

The following airspeeds are those which are significant to the safe operation

of the airplane. These figures are for standard airplanes flown at gross weight under

standard conditions at sea level.

Performance for a specific airplane may vary from published figures

depending upon the equipment installed, the condition of the engine, airplane and

equipment, atmospheric conditions and piloting technique.

(a) Best Rate of Climb Speed…........................................................110 KIAS

(b) Best Angle of Climb Speed........................................................... 81 KIAS

(c) Turbulent Air Operating Speed (See Subsection 2.1) .............. 133 KIAS

(d) Landing Final Approach Speed (Full Flaps).............................. 77 KIAS

(e) Maximum Demonstrated Crosswind Velocity ............................. 17 KTS

(f) Maximum Flaps Extended Speed

10°....................................................................................165 KIAS

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20°....................................................................................130 KIAS

Full Flaps (36°)................................................................116 KIAS

4.5f Ground Check Checklist (4.19.)

GROUND CHECK (4.19.)

CAUTION

Parking Brake............................................................................................. SET

Propeller Control…….......................................................... FULL INCREASE

Throttle..............................................................................................2000 RPM

Magnetos........................................................................... max. drop 175 RPM

- max. diff. 50 RPM

Gyro Suction............................................................................4.8 to 5.2 in. Hg.

NOTE

Ice protection equipment........................................ CHECK AS REQUIRED

Voltmeter ........................................................................................... CHECK

Ammeters ........................................................................................... CHECK

Oil Temperature ................................................................................ CHECK

Oil Pressure ....................................................................................... CHECK

Propeller Control............................................................... EXERCISE – then

FULL INCREASE

Fuel Flow ........................................................................................... CHECK

Throttle ............................................................................................ RETARD

Annunciator Panel ..............................................................PRESS-TO-TEST

Manifold Pressure Line ...................................................................... DRAIN

If flight into icing conditions (in visible moisture

below +5°C) is anticipated, conduct a preflight check

of the ice protection systems per Supplement No. 6 -

Ice Protection System.

Alternate air is unfiltered. Use of alternate air during ground or flight

operations when dust or other contaminants are present may result in

damage from particle ingestion.

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4.19 GROUND CHECK (4.5f)

Set the parking brake. The magnetos should be checked at 2000 rpm with the

propeller control set at full INCREASE. Drop off on either magneto should not

exceed 175 rpm and the difference between the magnetos should not exceed 50 rpm.

Operation on one magneto should not exceed 10 seconds. Conduct a preflight check

of the ice protection systems for proper operation.

Check the suction gauge; the indicator should read 4.8 to 5.2 in. Hg at 2000

rpm. Check that both red flow buttons are pulled in.

Check the voltmeter and ammeters for proper voltage and alternator outputs.

Check oil temperature and oil pressure. The temperature may be low for some time

if the engine is being run for the first time of the day.

The propeller control should be moved through its complete range to check for

proper operation and then placed in full INCREASE rpm for takeoff. Do not allow a

drop of more than 500 rpm during this check. In cold weather, the propeller control

should be cycled from high to low rpm at least three times before takeoff to make

sure that warm engine oil has circulated.

Check that the fuel flow gauge is functioning, then retard the throttle. Check

the annunciator panel lights with the press-to-test button.

Drain the manifold pressure line by running the engine at 1000 rpm and

depressing the drain valve, located on the left side of the control pedestal under the

instrument panel, for 5 seconds. Do not depress the valve when the manifold

pressure exceeds 25 inches Hg.

4.5g Before Takeoff Checklist (4.21)

BEFORE TAKEOFF (21.)

Battery Master Switch ...................................................................................ON

Alternators ...............................................................ON - CHECK AMMETERS

Pressurization Controls ................................................................................SET

Flight Instruments .................................................................................CHECK

Fuel Selector ..............................................................................PROPER TANK

Emergency (EMERG) Fuel Pump …………....................................................ON

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Engine Gauges ........................................................................................CHECK

Induction Air .......................................................................................PRIMARY

WARNING

NOTE

Pitot heat.................................................................................... AS REQUIRED

Stall warning heat...................................................................... AS REQUIRED

Wshld heat.................................................................................. AS REQUIRED

Prop heat.................................................................................... AS REQUIRED

Seat Backs................................................................................................ERECT

Seats………….........................................................adjusted& locked in position

Armrests ...............................................................................................STOWED

Mixture ............................................................................................ FULL RICH

Propeller Control ................................................................... FULL INCREASE

Belts/Harness .............................................................FASTENED/ADJUSTED

Empty Seats .............................................SEAT BELTS SNUGLY FASTENED

Flaps ..............................................................................................................SET

Trim ...............................................................................................................SET

Controls ......................................................................................................FREE

Door ....................................................................................................LATCHED

Air Conditioner ............................................................................................ OFF

Parking Brake ................................................................................. RELEASED

If flight into icing conditions (in visible moisture

below +5°C) is anticipated or encountered during

climb, cruise or descent, activate the aircraft ice

protection system, including the pitot heat, as

described in supplement No. 6 - Ice Protection

System.

Prolonged operation of the stall warning vane

heater in temperatures greater than 5ºC will

reduce the operational life of the stall warning

vane.

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4.21 BEFORE TAKEOFF (4.5g)

Ensure that the battery master and alternator switches are ON. Check that

the cabin pressurization controls are properly set. Check and set all of the flight

instruments as required. Check the fuel selector to make sure it is on the proper

tank. Ensure emergency (EMERG) fuel pump is ON. Check all engine gauges. The

induction air should be in the PRIMARY position.

Turn pitot, stall warning, windshield, and propeller heat ON if necessary.

Seats should be adjusted and locked in position. All seat backs should be erect

and armrests stowed.

The mixture control should be set to full RICH and propeller control should

be set to full INCREASE. Seat belts and shoulder harnesses should be fastened.

Fasten the seat belts snugly around the empty seats.

Set the flaps and trim. Ensure proper flight control movement and response.

The door should be properly latched and the door ajar annunciator light out. The air

conditioner must be OFF to ensure normal takeoff performance. Release the parking

brake.

4.5h Takeoff Checklist (4.23)

NOTE

NOTE

Takeoffs are normally made with full throttle.

However, under some off standard conditions, the

manifold pressure indication can exceed its

indicated limit at full throttle. Limit manifold

pressure to 42 in. Hg maximum. (See Section 7.)

During landing gear operation, it is normal for

the HYDRAULIC PUMP annunciator light to

illuminate until full system pressure is

restored.

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NORMAL TECHNIQUE (23a)

Flaps......................................................................................................0º to 10º

Trim............................................................................................................SET

Power................................................................................SET TO MAXIMUM

Liftoff.............................................................................................. 80-85 KIAS

Climb Speed................................................................................... 90-95 KIAS

Landing Gear (when straight ahead landing on runway not possible)…...UP

Flaps.................................................................................................RETRACT

0º FLAP TAKEOFF PERFORMANCE (4.23b)

Flaps............................................................................................................... 0º

Trim............................................................................................................SET

Brakes..,................................................................................................APPLY

Power................................................................................SET TO MAXIMUM

Brakes.............................................................................................. RELEASE

Liftoff................................................................................................... 78 KIAS

Obstacle Clearance Speed…............................................................... 91 KIAS

Landing Gear................................................................................................UP

SHORT FIELD TAKEOFF PERFORMANCE (23c.)

NOTE

Flaps............................................................................................................. 20º

Trim............................................................................................................SET

Brakes....................................................................................................APPLY

Power................................................................................SET TO MAXIMUM

Brakes............................................................................................. RELEASE

Liftoff................................................................................................... 69 KIAS

Obstacle Clearance Speed.................................................................. 80 KIAS

Landing Gear ...............................................................................................UP

Flaps................................................ RETRACT as speed builds thru 90 KIAS

Gear warning will sound when the landing gear is

retracted with the flaps extended more than 10º.

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4.23 TAKEOFF (see charts in Section 5) (4.5h)

NOTE

NOTE

Takeoffs are normally made with flaps 0º to 10º. For short field takeoffs or

takeoffs affected by soft runway conditions or obstacles, total distance can be

reduced appreciably by lowering the flaps to 20º.

4.23a Normal Technique (4.5h)

When the available runway length is well in excess of that required and

obstacle clearance is no factor, the normal takeoff technique may be used. The flaps

should be in the 0° to 10° position and the pitch trim set slightly aft of neutral. Align

the airplane with the runway, apply full power, and accelerate to 80-85 KIAS.

Apply back pressure to the control wheel to lift off at 80-85 KIAS, then control

pitch attitude as required to attain the desired climb speed of 90-95 KIAS. Retract

the landing gear when a straight-ahead landing on the runway is no longer possible.

Retract the flaps.

4.23b 0º Flaps Takeoff Performance (4.5h)

Retract the flaps in accordance with the Takeoff Ground Roll, 0º Flaps and

Takeoff Distance Over 50 Ft. Obstacle, 0º Flaps charts in Section 5. Set maximum

power before brake release and accelerate the airplane to 78 KIAS for liftoff. After

liftoff, adjust the airplane attitude as required to achieve the obstacle clearance

speed of 91 KIAS passing through 50 feet of altitude. Once immediate obstacles are

Takeoffs are normally made with full throttle.

However, under some off standard conditions,

the manifold pressure indication can exceed its

indicated limit at full throttle. Limit manifold

pressure to 42 in. Hg maximum. (See Section 7.)

During landing gear operation, it is normal for

the HYDRAULIC PUMP annunciator light to

illuminate until full system pressure is

restored.

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cleared, retract the landing gear and establish the desired enroute climb

configuration and speed.

4.23c Short Field Takeoff Performance (4.5h)

NOTE

For departure from short runways or runways with adjacent obstructions, a

short field takeoff technique with flaps set at 20º should be used in accordance with

the Takeoff Ground Roll, 20º Flaps and Takeoff Distance Over 50 Ft. Obstacle, 20º

Flaps charts. Maximum power is established before brake release and the airplane

is accelerated to 69 KIAS for liftoff. After liftoff, control the airplane attitude to

accelerate to 80 KIAS passing through the 50-foot obstacle height. Once clear of the

obstacle, retract the landing gear and accelerate through 90 KIAS while retracting

the flaps. Then establish the desired enroute climb configuration and speed.

4.5i Climb Checklist

MAXIMUM CONTINUOUS POWER CLIMB (4.25a)

Mixture........................................................................................ FULL RICH

Propeller Speed............................................................................... 2500 RPM

Manifold Pressure................................MAXIMUM CONTINUOUS POWER

Cylinder Head Temperature (CHT)...............................................500ºF MAX

Turbine Inlet Temperature (TIT)................................................1750ºF MAX

Oil Temperature............................................................................ 245ºF MAX

Best Angle of Climb (short duration only)..........................................81 KIAS

Best Rate of Climb............................................................................ 110 KIAS

Pressurization Controls............................................................................ SET

Emergency (EMERG) Fuel Pump........................................................ OFF at

safe altitude

Gear warning will sound when the landing gear is

retracted with the flaps extended more than 10º.

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NOTE

4.41 STALLS

The stall characteristics of the Malibu are conventional. An approaching stall

is indicated by a stall warning horn which is activated between five and ten knots

above stall speed. Mild airframe buffeting and pitching may also precede the stall.

The gross weight stalling speed with power off, landing gear extended, and

full flaps is 58 KIAS. With the landing gear retracted and flaps up, this speed is

increased to 69 KIAS. Loss of altitude during stalls can be as great as 700 feet,

depending on configuration and power.

NOTE

During preflight, the stall warning system should be checked by turning the

battery switch on and pressing the stall warning test switch to determine if the horn

is actuated.

2.14.4 SECTION 5 PERFORMANCE (Excerpts)

5.3 INTRODUCTION - PERFORMANCE AND FLIGHT PLANNING

The performance information presented in this section is based on measured

Flight Test Date corrected to I.C.A.O. standard day conditions and analytically

expanded for the various parameters of weight, altitude, temperatures, etc.

(Omitted)

An aircraft to have appropriate service following the procedure shown in the

performance table will recreate the performance.

(Omitted)

To obtain the performance shown in the performance table, do not forget to

follow the procedure written in the figure/table.

The stall warning system is inoperative with the

battery and alternator switches OFF.

For maximum engine life it is recommended to

transition to Cruise Climb once a safe altitude is

attained.

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WARNING

5.5 FLIGHT PLANNING EXAMPLE

(a) Aircraft Loading

The first step in planning the flight is to calculate the airplane

weight and center of gravity by utilizing the information provided by

Section 6 (Weigh and Balance) of this handbook.

(Omitted)

Make use of the Weight and Balance Loading Form (Figure 6-11)

and the C.G. Range and Weight graph (Figure 6-15) to determine the

total weight of the airplane and the center of gravity position.

(Omitted)

(b) Takeoff and Landing

(Omitted)

Apply the departure airport conditions and takeoff weight to the

appropriate “Takeoff Ground Roll and Takeoff Distance (Figures 5-13,

5-15 5-17 and 5-19)” to determine the length of runway necessary for

the takeoff and/or obstacle clearance.

(Omitted)

The conditions and calculation for flight plan are shown as follows;

The takeoff and landing distances required for the flight have fallen

below the available runway lengths.

Departure

airport

Designated

airport

(1) pressure altitude 1000 ft 1000 ft

(2) temperature 25℃ 25℃

(3) wind component 15 kts 10 kts

(4) length of runway 3400 ft 5000 ft

(5) required takeoff distance and

landing distance

2230 ft 1830 ft

Performance information derived by extrapolation

beyond the limits shown on the charts should not

be used for flight planning purposes

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2.14. 5 SECTION 6 WEIGHT AND BALANCE (Excerpts)

6.1 GENERAL

In order to achieve the performance and flying characteristics which are

designed into the airplane, it must be flown with the weight and center of gravity

(C.G.) position within the approved operating range (envelope). Although the

airplane offers flexibility of loading, it cannot be flown with the maximum number

of adult passengers, full fuel tanks and maximum baggage. With the flexibility

comes responsibility. The pilot must ensure that the airplane is loaded within the

loading envelope before he makes a takeoff.

Misloading carries consequences for any aircraft. An overloaded airplane will

not take off, climb or cruise as well as a properly loaded one.

The heavier the airplane is loaded, the less climb performance it will have.

Center of gravity is a determining factor in flight characteristics. If the C.G.

is too far forward in any airplane, it may be difficult to rotate for takeoff or landing.

If the C.G. is too far aft, the airplane may rotate prematurely on takeoff or tend to

nose-up during climb. Longitudinal stability will be reduced. This can lead to

inadvertent stalls and even spins; and spin recovery becomes more difficult as the

center of gravity moves aft of the approved limit.

A properly loaded airplane, however, will perform as intended. Before the

airplane is licensed, a basic empty weight and C.G. location is computed (basic

empty weight consists of the standard empty weight of the airplane plus the optional

equipment). Using the basic empty weight and C.G. location, the pilot can determine

the weight and C.G. position for the loaded airplane by computing the total weight

and moment and then determining whether they are within the approved envelope.

The basic empty weight and C.G. location are recorded in the Weight and

Balance Data Form and the Weight and Balance Record (Attachment 2 “Weight and

Balance Data Form”). The current values should always be used. Whenever new

equipment is added or any modification work is done, the mechanic responsible for

the work is required to compute a new basic empty weight and C.G. position and to

write these in the Aircraft Log Book and the Weight and Balance Record

(Attachment 2 “Weight and Balance Data Form”). The owner should make sure that

it is done.

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A weight and balance calculation is necessary in determining how much fuel

or baggage can be boarded so as to keep within allowable limits. Check calculations

prior to adding fuel to insure against improper loading.

The following pages are forms used in weighing an airplane in production and

in computing basic empty weight, C.G. position, and useful load. Note that the useful

load includes usable fuel, baggage, cargo and passengers. Following this is the

method for computing takeoff weight and C.G

6.7 GENERAL LOADING RECOMMENDATIONS

For all airplane configurations, it is the responsibility of the pilot in command

to make sure that the airplane always remains within the allowable weight vs.

center of gravity while in flight.

The following general loading recommendation is intended only as a guide.

The charts, graphs, instructions and plotter should be checked to assure that the

airplane is within the allowable weight vs. center of gravity envelope.

(a) Pilot Only

Load rear baggage compartment first.

(b) 2 Occupants - Pilot and passenger in front

Load rear baggage compartment first. Without aft baggage, fuel load

may be limited by forward envelope for some combinations of optional

equipment.

(c) 3 Occupants - 2 in front, 1 in rear

Baggage in nose may be limited by forward envelope.

(d) 4 Occupants - 2 in front, 2 in rear

Fuel may be limited for some combinations of optional equipment.

(e) 5 Occupants - 2 in front, 1 in middle, 2 in rear

Investigation is required to determine optimum baggage load.

(f) 6 Occupants - 2 in front, 2 in middle, 2 in rear

With six occupants fuel and/or baggage may be limited by envelope.

Load forward baggage compartment first.

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NOTE

Always load the fuel equally between the right and left tanks.

With takeoff loadings falling near the aft limit, it

is important to check anticipated landing loadings

since fuel burn could result in a final loading

outside of the approved envelope.

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2.15 Descriptions from the Maintenance Manual and Other Documents for

PA-46-350P

2.15.1 TIO-510-AE2A Operator’s Manual by the Engine Manufacturer

The following descriptions are included in TIO-510-AE2A Operator’s Manual by

the engine manufacturer:

(Excerpts)

With propeller in minimum pitch angle, set the engine to produce 50-65% power

as indicated by the manifold pressure gage. Mixture control should be in the full rich

position. At these settings, the ignition system and spark plugs must work harder

because of the greater pressure within the cylinders. Therefore, any weakness in the

ignition system will be more apparent. Mag checks at low power settings will only

indicate fuel-air distribution quality.

(Excerpts)

Correct power approximately 1% for each 10ºF variation in air temperature

from standard altitude temperature. Add correction for temperature below standard;

subtract correction for temperature above standard.

Figure 3-4 Sea Level/ Altitude Performance Curve

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2.15.2 Propeller Owner’s Manual (Manual No. 115N) by the Propeller Manufacturer

According to the Propeller Owner’s Manual (Manual No. 115N) by the Propeller

Manufacturer, adjusting the low pitch stop, not the propeller governor, in the checking

of the static RPM may cause a change to the blade angle at the low pitch stop and also

a change to rpm at that position.

(Excerpts)

(1) Set the brakes and chock the aircraft or tie aircraft down.

(2) Back the governor Maximum RPM Stop out one turn.

(3) Start the engine.

(4) Advance the propeller control lever to Max (max RPM), then retard the control

lever one inch (25.4 mm)

(5) SLOWLY advance the throttle to maximum manifold pressure.

(6) Slowly advance the propeller control lever until the engine speed stabilize.

(a) If engine speed stabilized at the maximum power static RPM specified by the

TC or STC holder, then the low pitch stop is set correctly.

(b) If engine speed stabilizes above or below the rated RPM, the low pitch stop may

require adjustment. Refer to the Maintenance Practices Section of this manual.

(61-00-15)

Furthermore, the manual includes the following warnings.

(Excerpts)

WARNING: SIGNIFICANT ADJUSTMENT OF THE LOW PITCH STOP TO

ACHIEVE THE SPECIFIED STATIC PRM MAY MASK AN ENGINE POWER

PROBLEM.

Refer to the following applicable procedure for accomplishing and adjustment

to the low pitch angle:

(Omitted)

Turning the low pitch stop screw one revolution equals 0.042 inch (1.06 mm) of

linear travel, and results in approximately 1.4 degree blade angle change. This blade

angle change results in an RPM increase/decrease of approximately 200 RPM.

2.16 Tests and Verifications Information

The teardown inspection of the engine, propeller, magnetos and air-conditioner

implemented to ascertain the conditions of the Aircraft at the time of the accident is

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outlined as follows:

(See Attachment 6 Teardown Inspection of Engine, Propeller and Others)

2.16.1 Teardown Inspection of the Engine

Regarding the engine of the Aircraft, a teardown inspection was carried out at

a facility of the engine manufacturer from January 12 to 13, 2016. The engine

manufacturer conducted the teardown inspection of the engine and the engine

accessories, and the propeller manufacturer did the teardown inspection of the

turbochargers. The whole of the engine showed the damage by the crash and the fire

after the crash. There was nothing found that would have preclude the engine from

making power prior to the crash.

As the result of the teardown inspection, no possibilities to lead to any

malfunction at the time of the crash were confirmed in the engine, the engine

accessories and the turbochargers.

2.16.2 Teardown Inspection of the Propeller

Regarding the propeller and the propeller governor, the propeller manufacturer

conducted the teardown inspection at the same time as the engine teardown inspection.

The propeller was found to be set at low-pitch angle or within that vicinity. One of the

blades had the scratch mark on the blade cord direction on the tip of the camber side,

and it means that the blade was rotating before the crash. The other propeller blade

was almost completely burnt down in the post-crash fire.

As the result of the teardown inspection, no discrepancy were noted that would

have prevented propeller operation prior to the crash including the propeller governor.

2.16.3 Teardown Inspection of the Magnetos

In order to investigate the ignition system of the engine, the teardown

inspection of the two magnetos as the major parts of the system was conducted on

August 2, 2016. The visual checks to the inside of the magnetos showed severe damage

on both magnetos due to the post-crash fire. Especially, the non-metal parts were lost

to fire, suffered severe damage and were charred. Due to these conditions, functional

inspections could not be carried out.

2.16.4 Teardown Inspection of the Air Conditioner

Regarding the operating status of the air-conditioning at the time of the takeoff,

the teardown inspection was carried out at the air-conditioning manufacturer on May

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9, 2016. It was not possible to determine whether the air-conditioner was operating or

not based on the residue status and the heated status due to the fire of the machine

oil of the refrigerator.

2.16.5 Verification of the Flight Route based on Images

2.16.5.1 The obtained visual materials

On the accident date, the flight situation was recorded in the multiple visual

images taken around Chofu Airport. The locations of the shooting of these visual

materials are indicated in Figure 2.16.5.1 and the visual materials are listed in Table

2.16.5.1.

Figure 2.16.5.1 Filming Locations

Table 2.16.5.1 List of Visual Materials

Filming location Type of visual images acronym

Observation deck located at the Terminal Camcorder: Handheld OBS

Baseball field E2 at Chofu Kichiatochi

Sports Park

Camcorder: Fixed E2

Softball field C at Osawa Sports Park Camcorder: Fixed SC

Softball field D at Osawa Sports Park Camcorder: Fixed SD

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Surveillance camera at Runway 35

Threshold

Fixed: Only visuals R35

Municipal Nishi-no-machi Soccer Field Camcorder: Handheld NS

Mainly in the cabin of the Aircraft Still images with GPS data SP

Hereinafter, use these acronyms to refer to the visual materials.

(1) Visual images of OBS

According to the visual images of OBS, the situation during the takeoff roll of

the Aircraft is as shown in Figure 2.16.5.1 (1).

Time The situation of the aircraft

10:57:35

10:57:36

10:57:37

10:57:38

Figure 2.16.5.1 (1) The Aircraft at the Takeoff Roll (OBS)

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(2) Visual images of R35

According to the visual images of R35, the situation during the climbing of the

Aircraft is as shown in Photo 2.16.5.1 (2), as a sequential photo.

Photo 2.16.5.1 (2) The Aircraft during the Climb after Takeoff (R35)

(3) Visual images of SC

According to the visual images of SC, the situation during the climbing of the

Aircraft is as shown in Photo 2.16.5.1 (3), as a sequential photo.

Photo 2.16.5.1 (3) -1 The Aircraft during the Flight (SC)

Photo 2.16.5.1 (3)-2 Condition of Black Smoke (SC)

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(4) Visual images of SD

According to the visual images of SD, the situation during the climbing of the

Aircraft is as shown in Photo 2.16.5.1 (4), as a sequential photo.

Photo 2.16.5.1 (4) The Aircraft during the Flight (SD)

2.16.5.2 Speed during the takeoff roll

The history of the speed during the takeoff roll was obtained from the images of

OBS, SP and E2. Because the shooting positions of these images were identified, the

time and the ground speed were calculated based on the comparison with the objects

on the ground seen in the images. The true air speed at no wind was first calculated

based on the ground speed obtained from the images and then the calibrated air speed

was estimated based on the calculated true air speed on the premise of the

temperature being 34ºC and the atmospheric pressure being 29.84 inHg. The speed

during the flight was estimated by the same method. This report, except as otherwise

noted, expresses the calibrated air speed as the speed. The obtained speed is shown in

Figure 2.16.5.2.

Figure 2.16.5.2 Estimated Ground Speed and Speed at the Time of Takeoff Roll

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2.16.5.3 The changes of speed and height after the takeoff

The angle of views, the lens distortions, the shooting directions and other

elements were measured based on the surveying from the shooting locations. The

north-south flight route was estimated based on the R35 images, and the position and

the height per time were estimated based on the images of SC and SD. The times of

the SC and SD images were synchronized after adding the sound propagation delay

caused by the distances between the crash site and the shooting locations onto the time

when the crash sound was recorded. Because the SD and R35 images contained images

recorded right after the takeoff, the synchronization was based on the estimated

position. The estimated speed from the SC and SD images are shown in Figure

2.16.5.3-1 and the estimated heights are shown in Figure 2.16.5.3-2. The estimated

speeds are vibrational, but this is due to the fluctuation of the reference point of the

airframe because the images of the Aircraft are blurred and small. The speeds (the

ground speed based on the smartphone GPS) recorded in the photos taken in the cabin

of the Aircraft are also shown in Figure 2.16.5.3-1.

Figure 2.16.5.3-1 Changes of the Ground Speed after Takeoff

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Figure 2.16.5.3-2 Changes of the Height after Takeoff

Integrating the results of these analyses mentioned above, the speed, height

and pitch angle 2.16of the Aircraft at the accident day were reconstructed against the

time are shown in Figure 2.16.5.3-3. The reconstructed speed and height against the

distances are shown in Figure 2.16.5.3-4.

Figure 2.16.5.3-3 Reconstructed Speed (CAS), Height and Pitch Attitude

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Figure 2.16.5.3-4 Reconstructed Speed (CAS) and Height

2.16.5.4 Estimated flight route

The estimated flight route is shown in Figure 2.16.5.4, together with the

outside photos taken in the cabin of the Aircraft with the time stamped on the photos.

Figure 2.16.5.4 Estimated Flight Route

(See Attachment 1 Photo Taken in the Cabin of the Aircraft (Right wing, flap and

others))

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2.16.5.5 Sound analysis during the before-takeoff check

The sound of the before-takeoff check was recorded in the E2 images.

Frequencies of these were analyzed using Sonic Visualizer*3, and the peak frequencies

are shown in Figure 2.16.5.5. The changes of the sound were matched with the sound

changes that are generated at the time of the check procedures defined in the flight

manual.

Figure 2.16.5.5 Sound during the Before-takeoff Check

2.16.5.6 Sound analysis during the flight

Sound during the flight were recorded in multiple images. Analyzing of these

with Sonic Visualizer, the sound of the propeller and the sound of the engine were

identified and the frequencies could be obtained. As an example, Figure 2.16.5.6 shows

the sound recorded in the SD images. The frequency of the propeller sound was

approximately 82 Hz. Because the propeller of the Aircraft has two blades, the rpm

was calculated as approximately 2,460 rpm and there were no significant fluctuations

observed. Moreover, no sound indicating any anomaly of the engine or other parts was

observed.

The sound during the flight contained noises, reflected and other sounds, and

this sound analysis did not take into account noise, reflected sound and the Doppler

effects by movement of the Aircraft.

Figure 2.16.5.6 Sound during the Flight 1

*3 Sound Analysis: Chris Cannam, Christian Landone, and Mark Sandler, Sonic Visualizer; An Open

Source Application for Viewing, Analyzing, and Annotating Music Audio Files, in Proceedings of the

ACM Multimedia 2010 International Conference.

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2.16.5.7 Readings of the instrument panel in the images taken in the cabin of the

Aircraft

(1) Information from manifold pressure meter and fuel flow meter

The parts of the manifold pressure meter and fuel flow meter could be seen in

the images taken in the cabin of the Aircraft. The enlarged images are shown in Figure

2.16.5.7(1).

Figure 2.16.5.7 (1) Manifold Pressure and Fuel Flow Meter Taken in the Cabin of

the Aircraft

Based on the readings of these meters, the manifold pressure was

approximately 40 inHg right after the start of the takeoff roll (10:57:29) and was

approximately 39 inHg while retracting the landing gear after the takeoff (10:57:53).

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(2) TIT gauge information

Figure 2.16.5.7 (2) TIT Gauge Taken in the Cabin of the Aircraft

The TIT (Turbine Inlet Temp) value could also be read in the same images that

show the manifold pressure meter and fuel flow meter as mentioned in (1). The

enlarged images are shown in Figure 2.16.4.7 (2).

Base on the readings of these, the TIT was approximately 980ºF during the

takeoff roll (10:57:29) and was 1,010ºF while retracting the landing gear after the

takeoff (10:57:53).

TIT is the temperature at the inlet of the turbocharger, which corresponds to

the exhaust temperature of the engine. The range from 1,200 to 1,750ºF is the normal

operating range of the Aircraft indicated as a green arc on the TIT meter. The flight

manual has a description of a maximum value of 1,750ºF as the operating limit, but

there is no setting for a minimum value. Upon a flight test of other aircraft of the same

type, TIT was approximately 1,400ºF at the maximum power.

(3) Landing gear warning light

At In the images at

10:57:29 as described in (1) and

(2) above, the landing gear

warning light, which shows that

the landing gear is being

retracted, did not illuminate, but

the images at 10:57:53 show the

landing gear warning light

illuminated.

(See Attachment 2 Photos Taken in the Cabin of the Aircraft (Instrument Panel)

Figure 2.16.5.7 (3) Status of Landing Gear

Warning Light(10:57:53)

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2.16.6 Analysis of the Past Flight

On July 22, 2015, the flight of the Aircraft was recorded by the camera installed

within the aerodrome. This flight was the last flight before the accident. From the

images, performance and flight control mainly at the time of takeoff roll was analyzed.

The flight record shows that the captain was piloting the Aircraft.

As shown in Photo 2.16.6, only the scene of passing the threshold of the runway

and climbing gradually was recorded, and based on this scene, the speed and height

at the time of passing through the threshold of the runway was calculated. Estimated

takeoff weight based on loaded fuel and the number of the persons on board obtained

from the submitted notification of the airport use, the meteorological data (outside

temperature, wind) and others are shown in Table 2.16.6-1

The position to start the takeoff roll of this flight is unknown, but if the starting

point is assumed to be about 10 m from the threshold of Runway 17, the threshold of

Runway 35 is the point about 790m from the starting point. Therefore, according to

the flight manual, the Aircraft should pass the threshold of Runway 35 at a height of

approximately 50 ft with a speed of slightly below 91 kt. However, based on the images,

the height and the speed were estimated to be 95 ft and 79 kt.

Photo 2.16.6 Images at the time of Takeoff Climbing on the Past Flight of the Aircraft

(July 22, 2015)

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Table 2.16.6-1 Specific Setting of the Flight on July 22, 2015 and Estimated Conditions

Estimated Takeoff Weight 1876 kg

Estimated Takeoff Time 14:30

Wind 14:00 210 degrees, 15 to 18 kt

15:00 210 degrees, 15 kt

Outside Temperature 14:00 33 ºC

15:00 32 ºC

Head Wind Component at 14:30 9 kt

Takeoff Ground Roll Distance 580 m

Passing Obstacle 50 ft distance 820 m

Table 2.16.6-2 Result of the Analysis of the Images Taken on July 22, 2015

Estimated height at the time of passing the threshold 95 ft

Estimated distance from point of starting the takeoff roll to the threshold 790 m

Estimated speed at the time of passing the threshold 79 kt

2.17 Information concerning the Operating and Maintenance Condition of

the Aircraft

According to the materials from the company which contracted for the

maintenance and administration of the Aircraft by the owner, the operating status and

others of the Aircraft was as follows:

(1) Annual flight time: according to the airworthiness inspection records, flight time

of the Aircraft over the last four years are as follows:

May 1, 2014 to May 1, 2015: 52 hours 23 minutes

May 2, 2013 to May 1, 2014: 32 hours 09 minutes

March 28, 2012 to May 2, 2013: 73 hours 29 minutes

February 10, 2011 to March 28, 2012: 106 hours 51 minutes

(2) Destinations of the flight: based on the restored flight records, which were

damaged due to the fire after the crash, the main destinations are as follows:

2015: Chofu local, Okadama, Fukushima, Shizuoka, Kagoshima

2014: Chofu local, Okadama, Fukushima, Shizuoka

2013: Chofu local, Sapporo, Sado, Oshima, Yao, Kouchi, Amami

2012: Chofu local, Okadama, Fukushima, Sado, Oshima, Oki

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2.18 Additional Information

2.18.1 Takeoff Ground Roll Distance and Takeoff Distance in Calculation

From the performance table (Attachment 3-1 to 3-4) of POH/AFM, the result of

calculating the Takeoff Ground Roll Distance*4 and Takeoff Distance*5are shown as

bellows. For the calculation, the temperature is set at 34 ºC as same as that at the

accident time, no wind, and the takeoff weight at 1,950 kg which is the maximum

takeoff weight.

(1) 0º flaps Takeoff

The Takeoff Ground Roll Distance: Approximately 2,230 ft (approximately 680 m)

The Takeoff Distance: Approximately 3,200 ft (approximately 976 m)

(2) Short Field Takeoff (at 20 º flap)

The Takeoff Ground Roll Distance: Approximately 1,730 ft (approximately 527 m)

The Takeoff Distance: Approximately 2,700 ft (approximately 823 m)

Based on Performance Table, it is possible to calculate a takeoff distance and a

takeoff ground roll distance to a takeoff weight up to the maximum takeoff weight.

However, if the weight exceeded the maximum takeoff weight, it is not possible to

calculate these distances. Moreover, even if it is assumed that it was flying at the

maximum takeoff weight, the takeoff ground roll distance based on the flight manual

of the Aircraft is shorter than the runway length of the airport (800 m), but the takeoff

distance was exceeding it.

According to the aircraft manufacturer, one should use the takeoff performance

table for 0º flap, following the 0º flap takeoff procedures, because the takeoff

performance using 10º flap is not different from the performance in the case of using 0º flap.

2.18.2 Regulations concerning Takeoff Weight

Pursuant to the Provision of Paragraph 3 of Article 10 of the Civil Aeronautics

Act “Airworthiness certification shall describe the categories of aircraft use and aircraft

operating limitations”, and pursuant to the provision of Paragraph 2 of Article 11 of the

Civil Aeronautics Act “No person may operate an aircraft beyond the categories of its

use or operating limitations as designated in the airworthiness certificate”. In addition,

pursuant to the provision of Paragraph 2 of Article 12-3 of Ordinance for Enforcement

of the Civil Aeronautics Act, aircraft operating limits shall be matters of limitations of

*4 “Takeoff Ground Roll Distance” means a horizontal distance from the standing point to start a

takeoff to the takeoff point.

*5 “Takeoff Distance” means a horizontal distance required to take-off and climb to a specified height

(50 ft for Category N) above the take-off surface

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aircraft, under the airworthiness examination guidelines (Ku-Ken No.381, enacted on

October 20, 1966) for aircraft of Aeroplane Normal N (hereinafter referred to as

“Category N”) as airworthiness categories which is corresponding to the accident

aircraft, maximum weight shall be matters of the limitations in flight manual. As

described in 2.14.1(2), maximum takeoff weight is described in flight manual of the

Aircraft as limitations.

As described in 2.6.1, the airworthiness category of the Aircraft fell under

Category N, and most privately owned airplanes flying over Japan fell under this

category. This airworthiness category means “an aircraft with a maximum certified

takeoff weight of 5,700 kg or less that is suited for normal flight (turns which do not

exceed 60 degrees in bank angle and stall (except a whip stall)) according to Annex 1

of the Ordinance for Enforcement of the Civil Aeronautics Act “Standards regarding

structures and performance to ensure the safety of aircraft and components”

(hereinafter referred to as the “Aircraft Standards”).

Most airplanes using the Airport fall under Category N, but the airworthiness

category of the aircraft used by air carriers who operate the regular service to and from

remote islands fall under Aeroplane Transport C (hereinafter referred to as “Category

C”). According to the Aircraft Standards, Category C means “multiple engine airplane

with a maximum takeoff weight of 8,618 kg or less that is suited for operations for air

transport services (limited to aircraft with 19 or less seats except that for a pilot).

Regarding takeoff performance of Category N in the airworthiness design

standards (hereinafter referred to as the “Airworthiness Standards”), which set the

requirements to show the compliance of the Aircraft Standards, it is required to set a

Takeoff Distance and describe it as performance data in a flight manual.

On the other hand, Category C is required to determine the acceleration-stop

distance, * 6 takeoff path* 7 and takeoff distance / takeoff roll distance as takeoff

performances and stipulate them, except takeoff path, as performance information in

the flight manual. Moreover, it is required that the flight manual of the category shall

contain the description of the maximum takeoff weight determined by applied with

acceleration-stop distance, takeoff distance, takeoff roll distance and the like, which is

determined depending on the runway length in use.

Moreover, as described in 2.8.1, the provision of Article 73-2 of the Civil

Aeronautics Law and Paragraph 1 of Article 164-14 of the Ordinance for Enforcement

*6 “Acceleration-stop distance” is the whole distance from a standing start point to the point where

the aircraft rejects a take-off and comes to full stop.

*7 “Take-off path” extends from a standing start point to a point at a height where the transition from

the take-off configuration to the enroute configuration must be completed.

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of the Civil Aeronautics Act obligate that the pilot in command shall be required to

confirm that the aircraft has no problems for flight regarding Takeoff weight, landing

weight, location of the C.G., and weight distribution.

In case of Aircrafts used for air transport services, air transport operators are

required to follow the Operation Manual which has obtained approval from the

Minister of Land, Infrastructure, Transport and Tourism, however, in “Detailed

Guideline of Examination of Operation Manuals (KuKoNo.78, enacted on January 28,

2000)”, a notice issued by the Director of Flight Operation Division, the Aviation Safety

and Security Department of Civil Aviation Bureau, which serves as the criteria for

granting approval of the manual prescribes that as the items for confirming takeoff

weight as stipulated in Article 164-14 of the Ordinance for Enforcement of the Civil

Aeronautics Act, which is required for preparing flight plans and determining whether

or not to take off as provided for in Article 73-2 of the Civil Aeronautics Act.

Chapter 3 Operation Manual Examination Standard (Part 2)

(an aircraft of maximum takeoff weight being 5,700 kg or less (except)

(Omitted)

2. Implementing method of a flight management

2-5 Flight management standard

As criteria to plan and change a flight, the following matters shall be

determined appropriately.

(1) Planning of Flight Plan and Decision whether to depart or not

f. Takeoff Weight, Landing Weight, C.G. position and Balance

(a) Takeoff weight and landing weight onto a dry runway which are

amended and calculated based on height of airport to be used, surrounding

obstacles, gradients of runway and others and weather condition shall be

complied with the following conditions.

Furthermore, for the conditions of wet, snow and ice, calculate to include

safety margins, appropriately. (If Flight Manual provides requirement,

follow the requirement.)

a. Takeoff weight and landing weight shall not exceed the maximum weight

as performance provided in Flight Manual.

b. Aircraft shall be the weight which requires takeoff distance to be less than

effective length of runway or landing strip (hereinafter referred to as “the

runway and others”).

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c. to d (Omitted)

(2) The C.G. position shall be within an allowable range.

The aircraft did not have any applicable Operation Manual since it was a private

aircraft, however, as described in 2.14.1 and 2.14.4, Section 2 “Limitations” of the flight

manual of the Aircraft has descriptions of “Note; See Section 5 “Performance” for

operating limitations of a maximum weight” and the paragraph of Section 5;

Performance has mentions with “To obtain the performance from Performance table, do

not forget to follow the procedure written in the Figure (Omitted) and “Apply the

departure airport conditions and takeoff weight to the appropriate “Takeoff Ground

Roll and Takeoff Distance (omitted)” to determine the length of runway necessary for

the takeoff and/or obstacle clearance.”

2.18.3 Statements concerning the Captain

(1) Statements of the instructor when the captain obtained a flight instructor

certification

According to the instructor, the captain had a good sense of control as a pilot.

Besides, the instructor remembers that the captain was honest, a good listener and

was quick to understand. However, the instructor also felt that the captain seemed

eager to obtain a flight instructor certification even though he did not have much flight

experience. Moreover, during the training, the instructor thought that as the captain

did not have enough experience as an air carrier pilot, he lacked the experience to

make judgments on his own in piloting. As the nature of the flight instructor

certification, how to teach is more important than the sense of piloting, and instructing

trainees while flying is quite difficult. The captain seemed to have hard time in

different situations to conduct training while controlling the flight from the unfamiliar

right seat.

The captain was a type of pilot who complied with checklists and manuals. The

instructor also stated that the captain surely must have prepared detailed flight plans.

(2) Statement of the person who has experience of flying the PA-46-350P

According to the person who has experience of flying the PA-46-350P, he saw

the captain landing an aircraft at the airport on a windy day, a few days before the

accident. He had conversations with the captain about his good piloting on such a day,

but the person told that he himself would not fly in such bad weather conditions.

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(3) Statement of the maintenance engineer who was a co-worker of the captain at

his previous company

According to the maintenance engineer who was a co-worker of the captain at

his previous company, he thought that the captain knew the weight and balance at the

time of the accident, as the captain used to calculate the weight of the aircraft for each

flight. On the other hand, he said that the captain was overconfident and might have

thought he would be able to fly even an over-weight aircraft.

2.18.4 Premature Nose-up at the Time of Takeoff

(1) “Airplane Flying Handbook” (FAA-H-8023, chapter 5 Takeoffs and Departure

Climbs) issued by FAA includes the following descriptions (Excerpts):

After rotation, the slightly nose-high pitch should be held until the airplane lifts

off. Rudder control should be used to maintain until the airplane lifts off. Rudder

control should be used to maintain the track of the airplane along the runway

centerline until any required crab angle in level flight is established. Forcing it into

the air by applying excessive back-elevator pressure would only result in an

excessively high-pitch attitude and may delay the takeoff. As discussed 5 earlier,

excessive and rapid changes in pitch attitude result in proportionate changes in the

effects of torque, thus making the airplane more difficult to control.

Although the airplane can be forced into the air, this is considered an unsafe

practice and should be avoided under normal circumstances. If the airplane is forced

to leave the ground by using too much back-elevator pressure before adequate flying

speed is attained, the wing’s AOA may become excessive, causing the airplane to settle

back to the runway or even to stall.

(Omitted)

Vx is the speed at which the airplane achieves the greatest gain in altitude for

a given distance over the ground. It is usually slightly less than Vy, which is the

greatest gain in altitude per unit of time. The specific speeds to be used for a given

airplane are stated in the FAA-approved AFM/POH. The pilot should be aware that,

in some airplanes, a deviation of 5 knots from the recommended speed may result in a

significant reduction in climb performance; therefore, the pilot must maintain precise

control of the airspeed to ensure the maneuver is executed safely and successfully.

(Omitted)

The pilot must always remember that an attempt to pull the airplane off the

ground prematurely, and to climb too steeply, may cause the airplane to settle back to

the runway or make contact with obstacles. Even if the airplane remains airborne,

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until the pilot reaches VX, the initial climb will remain flat, which diminishes the

pilot's ability to successfully perform the climb or clear obstacles.

The objective is to rotate to the appropriate pitch attitude at (or near) VX. The

pilot should be aware that some airplanes have a natural tendency to lift off well before

reaching VX. In these airplanes, it may be necessary to allow the airplane to lift-off in

ground effect and then reduce pitch attitude to level until the airplane accelerates to

VX with the wheels just clear of the runway surface.

(2) “Airplane Operation Textbook” supervised by the Civil Aviation Bureau (Japan

Civil Aviation Promotion Foundation, March 31, 2009, p90) has following

descriptions;

“Keeping the takeoff attitude established during the takeoff roll, lift the

airplane smoothly. If you force to takeoff by adding an excess back pressure before the

aircraft reaches the lift-off speed, it might result in high nose-up attitude to reduce the

speed and could result in a stall. The stall at the takeoff causes serious effects to the

airframe and human life.”

2.18.5 Flying on the “Backside”

For an aircraft in level flight at a constant speed, as the speed is slower, drag

(air resistance) becomes lower. However, the drag becomes higher below the speed (Vx)

at which the ratio between lift and drag becomes the maximum. Because thrust

corresponds to drag, the thrust needs to be increased as the speed decreases in order

to maintain level flight below Vx, that is, the engine power must be increased. The

high speed side of Vx is called “the front side” and the low speed side is called “the

backside”.

During the flight on the front side, pitching up the nose from the level flight

while keeping engine power constant causes a decrease of the speed and the aircraft

starts to climb. This means that a part of the excess engine power due to the decrease

of the speed was directed to cause climbing.

On the other hand, if a similar maneuver is carried out during the flight on the

backside, the aircraft descends at the same time as the speed is decreased. This is

because the power becomes insufficient to maintain a level flight due to the decreased

speed. This kind of characteristics is sometimes called “the backside characteristics”.

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In pilot training, the subject called “slow flight” is intended to have trainees

obtain the maneuvering sense

during the flight on the

backside. However, if a speed is

decreased unintentionally to the

level entering into the backside,

pitching up to climb may result

in rapid loss of speed and height.

In order to escape from the

backside, it is necessary to lower

the nose and increase the speed,

but in order to avoid losing the

height, it is also necessary to

increase the power at the same

time. Figure 2.18.5 Speed and Thrust at Backside

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2.18.6 Safety Measures upon Landing and Takeoff of Aircraft at the Airport

Facilities

Airport facilities must be constructed based on the provision of Article 79 of the

Ordinance for Enforcement of the Civil Aeronautics Act (hereinafter referred to as the

“construction standards”). The Civil Aviation Bureau provides matters for

determining the position, forms, strength and others required for facilities from the

perspective of securing functionalities, safety, economy and others when designing

airport facilities, and also prepares a manual of the standards for constructing airport

facilities (hereinafter referred to as the “manual of the construction standards”) as an

explanatory document concerning the construction standards with the aim of

achieving the efficiency and improvement in the facility design and construction.

(1) Overrun areas

According to the manual of construction standards, a runway is defined as a

rectangular part provided for departure and arrival of aircraft and overrun areas are

stipulated as facilities provided on both ends of the runway in preparation for the case

where aircraft fails to stop within the runway or other cases.

(2) International standards and current status in Japan regarding runway end

safety areas

In order to mitigate the damage to the aircraft in the event of “Overrun”, which

means that an aircraft stops after passing the end of the runway upon landing or

taking off, or “Undershoot”, which means that an aircraft lands before the runway,

runway end safety areas need to be established at both ends of the landing strip,

including overrun areas, based on Annex 14 to the Convention on International Civil

Aviation (hereinafter referred to as “4 the Convention”).

At most airports in Japan, 40 m-long runway end safety areas have been

secured. The provisions of Annex 14, which were in a form of a recommendation, were

revised in 1999 and became the Standard, which requires extension of the length of

runway end safety areas to 90 m for runways of 1,200 m or more in length or

instrument landing runways. *8

Therefore, the Civil Aviation Bureau revised the manual of construction

standards in April 2013 requiring the improvement of runway end safety areas based

on Annex 14, in principle. Regarding airports where the length and width of runway

*8 “Instrument landing runway” is a runway for landing of aircraft using guidance (a flight completely

depending on instruments in measuring attitude, altitude, location and coourse of the aircraft).

Instrument approach is carried out completely depending on instruments such as NDB, VOR, DME,

RNAV, ILS and others, and has two types: precision approach and non-precision approach.

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end safety areas are not secured, the Bureau decided to evaluate the current status

concerning the factors that may lead to accidents and an extent of the damage in the

event of an accident, and take measures such as introducing an arresting system*9

and securing runway end safe areas, if the effects of those factors are evaluated to be

significant.

(3) Overrun areas and runway end safety areas at the Airport

The Airport has the overrun areas of 60 m in length and 30 m in width on both

ends of the runway based on the manual of the construction standards. The strength

of the pavement at the overrun areas is weaker than that on the runway.

According to the manual of construction standards revised in April 2013,

runway end safety areas that are 90 m or more in length (or preferably 120 m or more,

if possible) and 60 m in width with 5% or less gradient in vertical and horizontal

directions are required to be installed at the end of the overrun areas at an airport for

instrument landing even if it has an 800 m-long runway. According to the aerodrome

provider and administrator (Tokyo Metropolitan Government), the land for runway

end safety areas was reserved as of the time of the accident, but the current status

evaluation (interim) as runway end safety areas was incomplete. The current status

evaluation was completed in March 2017 and it was concluded that if the land

corresponding to runway end safety areas is maintained and operated, there will only

be little damage to the aircraft in the event of an accident and no factor leading to

occurrence of an accident is found.

Figure 2.18.6-1 Runway, Overrun Area and Runway End Safety Area of the Airport

Figure2.18.6-2 Legend of Overrun area markings

*9 “Arresting system” means a mechanism to reduce the speed of an overrunning aircraft and mitigate

damage thereto. It is an alternative measure when the length or width of a runway end safety area

cannot be secured. However, the arresting system is a countermeasure for overrun, not for undershoot.

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(4) Measures to make maximum use of a runway length

In Japan, there are cases where entrance taxiways are connected to extensions

of runway as seen in Osaka International Airport and others, and enable a maximum

use of runway length in comparison to typical airports where entrance taxiways are

vertically connected to runway thresholds.

Figure 2.18.6-3 Case of Osaka International Airport

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Figure 2.18.6-4 Case of Kansai International Airport

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2.18.7 Status of the Maintenance Management of the Aircraft

According to the person in charge at the company which had provided the

service to maintenance and management of the Aircraft, he checked the records of the

maintenance work whenever a maintenance engineer provides maintenance services.

As far as the company found during checks on the Airplane on the ground, there were

no anomalies. The company did not think that the Aircraft had conducted flights with

malfunctions.

As for the engine power data, when the RPM and manifold pressure show

correct values during a test flight, it is determined that there is no problem. Regarding

TIT, if other data show correct values, it is tend to be judged as no problem. Regarding

trouble shooting, the company told that they might make inquiries directly to the

manufacturer or via an agent when they could not fix malfunctions, but they seldom

closely examined the data obtained through test flights. They were not aware of low

values of TIT gauge as there were no reports about it from the maintenance engineer.

2.18.8 Status of Implementation of the Airworthiness Directive

The contents and the implementation status of the Airworthiness Directive

regarding TIT gauge corresponding to the Aircraft were described as follows:

As the purpose to prevent malfunction leading to a loss of aircraft control

because a generator damaged due to inappropriate calibration of TIT gauge indicating

system and probe defects, the Airworthiness Directive TCD-5111-2000 (hereinafter

referred to as “the TCD”) was issued on 2000, which was fully revised to the

Airworthiness Directive TCD TCD-5111-2011 (hereinafter referred to as “the revised

TCD”). Implementing status of the revised TCD on April 25, 2014 at the total flight

time of 2,208 hours and 38 minutes of the Aircraft, the probe was replaced. According

to the maintenance records following the TCD and the revised TCD, the probe was

inspected for three times and the probe was replaced for six times. In addition, the

TCD and the revised TCD were issued by the Civil Aviation Bureau based on the

relevant AD 2011-06-10 and AD 99-15-04 R1, incorporated with the Aircraft

Manufacturer’s Service Bulletin No.995 and the latest revision to it.

2.18.9 International Standards, Oversea Regulations and Reference Cases

(1) Standards and Recommended Practices of Parts II (International General

Aviation – Aeroplanes) of Annex 6 to the Convention on International Civil Aviation

The Standards and Recommended Practices of Parts II (International General

Aviation – Aeroplanes) of Annex 6 to the Convention on International Civil Aviation

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have the following requirement as international standard regarding the confirmation

by the pilot-in-command before departure.

2.3.1.3 The pilot-in-command shall determine that aeroplane performance will

permit the take-off and departure to be carried out safely.

(2) Regulations set forth by the Federal Aviation Administration of United States

of America

The Federal Aviation Regulations have the following requirements regarding

the confirmation by pilot in command before departure:

91.103 Preflight action

Each pilot in command shall, before beginning a flight, become familiar with all

available information concerning that flight. This information must include-

(Omitted)

b) For any flight, runway lengths at airports of intended use, and the

following takeoff and landing distance information:

For Civil Aircraft for which an approved Airplane or Rotorcraft flight

manual containing takeoff and landing distance data is required, the takeoff

and landing distance data is required, the takeoff and landing distance data

contained therein;

(3) Regulations of the Civil Aviation Authority (United Kingdom)

The Air Navigation Order 2009 of United Kingdom stipulates the following

requirements: regarding the confirmation by pilot in command before departure:

PART 10

Duties of commander

(Omitted)

Commander to be satisfied that flight can be safely completed

87. The commander of a flying machine must, before take-off, take all reasonable

steps so as to be satisfied that it is capable of safely taking off, reaching and

maintaining a safe height and making a safe landing at the place of intended

destination having regard to –

(a) the performance of the flying machine in the conditions to be expected on

the intended flight; and

(b) any obstruction at the places of departure and intended destination and

on the intended route.

In addition, “SAFETYSENSE LEAFLET 7c AEROPLANE PERFORMANCE”

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January 2013 issued by the Civil Aviation Authority concerning small aeroplane

performance has the following descriptions (abstract):

(a) Introduction

The pilot in command has a legal obligation under EU Part-NCO and

Article 87 of the Air Navigation Order 2009, which require the pilot to check

that the aeroplane will have adequate performance for the proposed flight.

(b) TAKEOFF-POINTS TO NOTE

a Decision point: You should work out the runway point at which you can

stop the aeroplane in the event of engine or other malfunctions, e.g. low

engine rpm, loss of airspeed, lack of acceleration or dragging brakes. Do

NOT mentally programme yourself in a GO-mode to the exclusion of all

else.

b Use of available length: Make use of the full length of the runway; there

is no point in turning a good length runway into a short one by doing an

‘intersection’ takeoff.

(4) Report* 10 issued by the Australian Transport Safety Bureau (hereinafter

referred to as the ATSB)

According to the report issued by the ATSB, among events that occurred from

January 1, 2000 to January 1, 2010, and are reported to the ATSB, 242 cases were

partial power loss of small aeroplanes after takeoff (including 9 death cases) and 75

cases were engine malfunction after takeoff (no death case).

Major contents of “Avoidable Accidents No.3 Managing partial power loss after

takeoff in single-engine aircraft” are outlined as below.

This ATSB booklet aims to increase awareness among flying instructors and

pilots of the issues relating to partial power loss after takeoff in single-engine aircraft.

Most fatal and serious injury accidents resulting from partial power loss after

takeoff are avoidable. This booklet will show that you can prevent or significantly

minimise the risk of bodily harm following a partial or complete engine power loss

after takeoff by using the strategies below:

a Pre-flight decision making and planning for emergencies and abnormal

situations for the particular aerodrome

b producing a thorough pre-flight and engine ground run to reduce the risk

of a partial power loss occurring

c taking positive action and maintaining aircraft control either when

*10 Details of the report concerning “managing partial power loss after take-off in single-engine

aircraft” issued, studied and investigated by the Australian Transport Safety Bureau are publicized on

the ATSB website (http://www.atsb.gov.au).

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turning back to the aerodrome or conducting a forced landing until on the

ground, while being aware of flare energy and aircraft stall speeds.

Examples of the causes of engine power loss include, but are not limited to:

a mechanical discontinuities within the engine

b restricted fuel or air flow or limited combustion in the engine, often due

to fuel starvation, exhaustion or spark plug fouling

c mechanical blockage in the engine setting controls, such as a stuck or

severed throttle cable.

A partial engine power loss presents a more complex scenario to the pilot than

a complete engine power loss. Pilots have been trained to deal with a complete power

loss scenario with a set of basic check and procedures before first solo flight. (Omitted)

in a partial power loss, pilots are faced with making a difficult decision whether to

continue flight or to conduct an immediate forced landing.

By extending already established procedures dealing with total power loss to a

partial engine power loss scenario, this report will present the different options to

consider during your pre-flight planning.

a pre-flight planning (which focuses on preparing for loss of power)

b avoiding a partial lower loss after takeoff

- operations on the ground (preventing loss of power)

- the pre-takeoff self-briefing

- on takeoff checks and rejecting the takeoff.

c managing a partial power loss after takeoff (planning considerations and

maintaining control)

- forced or precautionary landing (on or beyond the air field)

- turning back towards the departure aerodrome.

Summary

(a) Pre-flight checks prevent partial power loss

ATSB occurrence statistics indicate that many partial power losses could

have been prevented by thorough pre-flight checks. Some conditions reported as

causing partial power loss after takeoff are fuel starvation, spark plug fouling,

carburetor icing and pre-ignition conditions. In many cases, these conditions

may have been identified throughout the pre-takeoff and on-takeoff check

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phases of the flight sequence.

(b) Pre-flight planning and pre-takeoff briefings

Even if a partial power loss does occur after takeoff, considering actions

to take following a partial power loss after takeoff during the process of planning

and the pre-flight safety brief gives pilots a much better chance of maintaining

control of the aircraft, and helps the pilot respond immediately and stay ahead

of the aircraft. Considerations include planning for rejecting a takeoff, landing

immediately within the aerodrome, landing beyond the aerodrome, and

conducting a turnback towards the aerodrome.

(c) Stay in control

If nothing else, maintain glidespeed and plan a maximum bank angle

against your personal minimums, which you will not exceed if a turnback is an

option. Be prepared to re-assess the situation throughout any maneuver.

(5) Leaflets by the Civil Aviation Authority of New Zealand (hereinafter referred to

as the “CAA”)

According to the materials of the CAA, during the time from January 1, 1995 to

December 31, 2012, the CAA received reports of 59 cases of partial power loss

(including 3 deaths and 12 persons with serious injuries). Among these, a little over

30% occurred during the time of taking off and climbing. Categorizing the causes of

occurrences, mechanical malfunction was 55%, piloting 11%, icing 9%, fuel related 6%

and other causes or unknown 19%.

Major contents of the VECTOR May/June 2013 Partial Power issued by the CAA

are outlined as follows, indicating the countermeasures against engine power loss:

(a) Preflight Planning

By considering the many factors involved in the takeoff, such as wind

strength and direction, runway direction, terrain and obstacles, and landing

options on and off the airfield, you will reduce the mental workload required to

handle a loss of power. This can also help you with decision making under stress

or a high workload in an emergency.

Getting this plan together before you leave, will give you the confidence

to carry out timely and positive actions if required.

(b) Preflight Checks and Inspection

The preflight inspection is a vital action for any flight and can reduce the

likelihood of a partial power loss occurring after takeoff.

(Omitted)

Ensure the engine starts easily and runs smoothly, and allow an

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adequate warm-up time.

Conducting a thorough engine run-up is an important step.

Testing fuel flow from the selected tank (fullest or takeoff tank), checking

for correct operation of the carburetor heat control, and checking and comparing

individual magnetos for a specified RPM drop range is vital. Engine oil

temperature and pressures, fuel pressure and other engine or systems gauge

indications should be within accepted aircraft operating limitations.

Allow plenty of time to conduct the engine run-up check to help show any

abnormalities with both the engine and fuel system, and never attempt to take

off when the engine continues to misfire or is running rough.

(c) Fuel

Fuel starvation, exhaustion, or contamination, also rate highly as causes

of partial power often leading to total power loss.

(Omitted)

(d) Induction Icing

(Omitted)

(e) Pre-flight Self-briefing

All single-engine aircraft pilots, just like multi-engine aircraft pilots,

should ‘self-brief’ before each and every takeoff. It helps you keep ahead of the

aircraft, and keep control.

This brief is generally conducted once all engine and systems checks are

complete, just prior to the holding point for takeoff. It serves as a reminder of

your planned actions in the event of an emergency.

Here is an example of a self-brief:

a Engine failure before rotation point, I will abort the takeoff, close the

throttle, and stop on the remaining runway.

b Engine failure after rotate, runway remaining, I will lower the nose,

close the throttle, land in the remaining runway available.

c Engine failure in initial climb, I will lower the nose, close the throttle,

select the best option, and execute trouble checks if time permits.

On the takeoff run, we wisely choose to use the full length of the runway

available, and on application of full power we check the static RPM to confirm

engine performance.

With the brakes off we check the acceleration of the aircraft, and the

performance of the engine for any signs of power loss and/or rough running.

After rotation and in the initial climb, any partial engine power loss that

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degrades performance to the extent that you cannot maintain height can be

treated as a complete engine failure with a potentially extended glide distance.

At this point, you might hear your instructor reminding you to, “lower

the nose to the gliding attitude, maintain speed, carry out trouble checks if you

have time, and fly the aircraft to a landing.”

(Omitted)

At a reasonable height, and with power that is sufficient to maintain

height, a turn back to the recently departed runway may be an option, but it

has a number of considerations attached. The overriding thought is that the

engine could fail at any time.

Accidents occur when control is lost, especially when the pilot attempts

to turn back to the runway at low level and low speed, or does not maintain

control in the glide.

2.18.10 Actions Taken by the Civil Aviation Bureau (Japan) after the Accident

The Civil Aviation Bureau took following measures after the occurrence of the

accident.

Monday, July 27, 2015:

The Civil Aviation Bureau published a notice to the operators of small

aircrafts requesting them to take every possible measure of operation for

reliable implementation of checking and maintenance and to comply with

regulations and procedures.

Wednesday, August 26 to Thursday, August 27, 2015:

In order to confirm the implementation status of the notice issued on July

27, officers of the Civil Aviation Bureau conducted a temporary safety audit

targeting nine operators based in Chofu City.

Friday, August 28, 2015:

The aerodrome provider and administrator (Tokyo Metropolitan

Government) and the Civil Aviation Bureau established a conference

regarding aircraft safety at Chofu Airport and held meetings. It was

decided that information regarding the result of the safety audit and the

studies of safety actions should be shared and safety measures should be

discussed and carried out in full cooperation.

Tuesday, September 1 and Wednesday, September 2, 2015:

Officers of the Civil Aviation Bureau confirmed the status of the Airport

after resuming of commercial flight operation.

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Thursday, November 26, 2016

Officers of the Civil Aviation Bureau held seminars to promote safety in

maintenance work for small aircrafts at Yao Airport. Following this,

seminars were also held in Chofu, Sendai, Nagoya, Tokyo and other places.

2.18.11 Actions Taken by the Aerodrome Provider and Administrator (Tokyo

Metropolitan Government) after the Accident

(1) On Tuesday, August 18, 2015, the aerodrome provider and administrator

(Tokyo Metropolitan Government) held a briefing targeting nearby residents and

provided information of the Aircraft, and explained the developments leading to the

accident, the actions it has taken as the aerodrome provider and administrator and

the future actions, as described below:

1) Immediate response

a. Regarding commercial aircraft, the operation will be resumed after a safety

check to confirm the safety of the aircraft to be conducted by maintenance

engineers with a national qualification, and after a special safety seminar to be

provided by lecturers from outside.

Similar safety checks will also be implemented once every 3 months this

fiscal year, and after the resuming the operation, a safety seminar will be

implemented periodically.

b. Taking special consideration of the fact that it was a privately owned

aircraft that caused the accident, owners are continuously requested to refrain

from flying their privately owned aircraft until the cause of the accident is

cleared and safety measures are put in place.

2) Future actions

Learning a lesson from the accident, the aerodrome provider and

administrator will take further safety actions to prevent reoccurrences, while

listening to the opinions of local citizens.

a. Verify whether the use of aircraft at the Airport has been appropriate, or

not

b. Promote discussions between local cities and the aerodrome provider and

administrator (Tokyo Metropolitan Government) regarding the strengthening

of safety actions and further improvement of the administrative management

of the Airport.

(2) The aerodrome provider and administrator (Tokyo Metropolitan Government)

held briefing sessions targeting nearby residents on June 16 (Thursday), 17 (Friday),

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and 20 (Monday), 2016 and explained future actions to be implemented as follows.

1) A captain, a maintenance engineer, an operation administrator and others

are to be obliged to participate in safety seminars. At the same time, the captain

of a privately owned aircraft will be required to thoroughly carry out a pre-

departure check and make a report to the administration office. Mechanical

engineers working at the Airport will be obliged to participate in lectures targeting

engineers held by the national government, and checking and maintenance of

aircraft must be implemented by those engineers.

2) In order to enhance the system of responsibility in the event of an accident

or other emergency, it will be required to appoint a person in charge of emergency

response for each privately owned aircraft. Persons thus appointed should hold

liaison conferences regularly and bear obligation to provide compensation and

apology to victims quickly in the event of an accident. Besides, owners of privately

owned aircrafts will be obliged to purchase aircraft third party liability insurance.

Moreover, the aerodrome provider and administrator (Tokyo Metropolitan

Government) will establish a consultation counter to meticulously deal with

requests and consultations from victims, and when relief measures are not taken

promptly, the aerodrome provider and administrator (Tokyo Metropolitan

Government) will responsibly provide aid to victims promptly such as by securing

temporary housing or removing damaged houses.

3) In order to ensure reasonableness of flights of privately owned aircraft, efforts

will be continued to thoroughly disseminate the fact that sightseeing flights are

not allowed under the agreements, concluded with the nearby local cities and other

new measures to prevent sightseeing flights will also be taken. If any sightseeing

flight is found by the privately owned aircraft, the aircraft will not be allowed to

use the Airport.

4) The Tokyo Metropolitan Government has worked to promote the positive

moving out of privately owned aircraft from the Airport based on the agreement

concluded with the nearby local cities in 1997; however, seriously considering

requests of three local cities and the regional conference of three municipalities

around the Airport asking for the elimination of privately owned aircraft from the

Airport, the Metropolitan Government will endeavor to reduce the number of

privately owned aircraft to the extent possible. For this purpose, the Metropolitan

Government will carry out in-detail investigation on the actual status of privately

owned aircraft at the Airport and places to relocate them, and promote concrete

actions therefor.

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5) Excluding the minimal operations of privately owned aircraft for the purpose

of taking an airworthiness inspection defined by laws and regulations or keeping

one’s piloting skill, the Metropolitan Government will continue to request owners

to refrain from flying their privately owned aircraft. Upon the first flight, an

examinee of a specific pilot competency assessment will be on board to confirm the

competence of the pilot.

3. ANALYSIS

3.1 Qualifications of Personnel

The Captain held both a valid airman competence certificate and a valid

aviation medical certificate.

3.2 Airworthiness Certificate of the Aircraft

The Aircraft had a valid airworthiness certificate and had been maintained and

inspected as prescribed.

3.3 Takeoff Weight and Balance of the Aircraft

As described in 2.6.4, it is highly probable that when the accident occurred, the

weight of the Aircraft is estimated to have been approximately 2,008 kg (approximately

4,427.5 lb) and that the Aircraft was heavier approximately 58 kg than the maximum

takeoff weight, which was 1,950 kg. Moreover, it is highly probable that the position of

the C.G. was +146.0 to +146.5 in. aft of the reference line, and it was close to the aft limit

at the time of the maximum takeoff weight.

The takeoff weight and the position of the C.G. were analyzed as below.

3.3.1 Confirmation of Weight and the position of C.G. before Departure

As described in 2.8.1, the pilot in command was required to confirm the takeoff

weight and the position of the C.G. before departure pursuant to the provisions of

Article 73-2 of the Civil Aeronautics Act, and Article 164-14 of the Ordinance for

Enforcement of the Civil Aeronautics Act.

As described in 2.6.4, the documents and other materials, which shows the

captain calculated the weight and the position of the C.G. of the Aircraft before

departure, have not been discovered. As described in 2.1.2 (2) and (3), Passengers B

and C stated that they were not asked about their weights by the Captain or Passenger

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A before the flight.

Judging from the facts above, it is probable that the weight and the position of

the C.G. of the Aircraft were not confirmed sufficiently by the Captain before departure.

3.3.2 The Captain’s Recognition about Takeoff Weight

As described in 2.6.4 and as analyzed in 3.3, it is highly probable that the takeoff

weight at the time of the accident exceeded the maximum takeoff weight.

In general, without accurate calculation, it is probable that pilots can estimate

the weight of their accustomed aircraft from their experience and expect whether the

weight will completely exceed or is likely to exceed the maximum takeoff weight based

on the amount of remaining fuel, number of persons on board and other conditions.

As for the situation of the Aircraft weight when the accident occurred, as

described in 2.6.4, it is highly probable that the weight of the loaded fuel was

approximately 278 kg, approximately 83 % of the total fuel capacity (333 kg). The most

recent flight of the Aircraft was also made by the Captain. Therefore, it is probable

that he could estimate the fuel weight situation.

It was not possible to determine whether the Captain recognized the weight

exceeding the maximum takeoff weight prior to the Accident flight because the

Captain died, but it is somewhat likely that he had insufficient understanding of the

risks of making flights under such situation and insufficient safety awareness of

observing laws, regulations and provisions.

3.3.3 Effects of Weight and Balance of the Aircraft

As described in 2.8.2 and 2.14.4(1), if an aircraft exceeding its maximum takeoff

weight engages in flight, the aircraft has reduced takeoff and climb performance and

cannot take off, climb or cruise as usual.

When the position of the C.G. is close to the aft limit, in the state that nose up

is easily to occur, the following effects will be occurred: excessive nose-up attitude at

the time of takeoff and during climb, or decreased controllability, stability or flight

performance during a low-speed flight which may lead to an unexpected stall.

Therefore, it is necessary to maneuver and control aircraft carefully.

3.3.4 Deviation from Allowable Ranges for Weight and the Position of the C.G..

The deviation from allowable ranges for the weight and the position of the C.G.

is the excess of operating limitations specified in airworthiness certificates. In addition,

as described in 3.3.3, the deviation causes takeoff performance or climb performance

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to decrease, and controllability, stability or flight performance to decrease, which may

lead to a stall during low-speed flight. Therefore flights must never be made under

such conditions.

In this accident, it is highly probable that the excess of the Aircraft weight

deteriorated the takeoff and climb performance of the Aircraft, and in the state that

nose-up was easily to occur because of the position of the C.G. close to the aft limit was

leading to the excessive nose-up attitude at the time of takeoff and climb, and

decreased controllability, stability and flight performance during a low-speed flight,

which brought the Aircraft to a condition easy to stall. It is highly probable that these

were the factors that led the Aircraft to take off at low speed, have the excessive nose-

up attitude, and enter stall.

The flight manual described in 2.14.5 shows the importance of the weight and

the position of the C.G. to be within allowable ranges. The flight manual also describes

that calculation of the weight and the position of the C.G. of aircraft are required in

order to keep them within allowable ranges and determine allowable loading capacity

of fuel or baggage. The flight manual further states that calculation results have to be

inspected prior to refueling in order to prevent inadequate loading status. When

planning flights, captains must accurately identify the weight of persons, baggage on

board and fuel, calculate the weight and the position of the C.G. based on these factors

by using calculation documents or other materials, and be sure to confirm that the

weight and the position of the C.G. are within allowable ranges. When deviations from

allowable ranges occur, captains are required to restrict persons and baggage to be

boarded or reduce the amount of fuel so that maintain the weight and the baggage of

the C.G. are maintained within allowable ranges before making flights.

In addition, it is probable that in some cases, it may be practically difficult to

reduce the amount of fuel once loaded. Therefore the amount of fuel to be loaded should

be carefully examined when there is a possibility that the weight of the aircraft is

likely to exceed the maximum takeoff weight or the weight restricted from the airplane

performance.

3.4 Flight of the Aircraft at the Time of the Accident

The flight route estimated from the flight history described in 2.1.1 and images

and other materials described in 2.16.5 was analyzed.

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3.4.1 Configuration of the Aircraft and Weather Conditions at the Time of the

Accident

(1) At the Time of Starting a Takeoff Roll

Judging from the statements of Passenger A described in 2.1.2 (1), it is probable

that the Aircraft underwent standing takeoff for which the pilot stopped the Aircraft

once on the runway, increased engine rotation sufficiently, and started a takeoff roll.

(2) Flaps of the Aircraft

Judging from the photos (such as the right wing) shown in Attachments 1-1 and

1-2, which were taken from inside the Aircraft, it is highly probable that the flap

setting were 10°.

(3) Landing Gear of the Aircraft

As described in 2.16.5.7 (3), the photo taken from inside the Aircraft at 10:57:53

shows that the landing gear warning was lit. As described in 2.7.2 (1), judging from

the fact that the landing gear warning is lit when retractable landing gear is being

retracted, it is highly probable that the gear of the Aircraft was being retracted at

10:57:53.

(4) Weather Conditions

From the aeronautical weather observations for Chofu Airport described in 2.9.2,

it is highly probable that temperature was 34°C.

As for the wind situation, judging from the aeronautical observation values for

Chofu Airport described in 2.9.2 (1) and the wind data around the time of the accident

as described in 2.9.2 (2) and Attachment 4, it is highly probable that it was not a tail

wind condition which affected the flight performance of the Aircraft. In addition,

judging from the black smoke rose almost vertically as shown on Photo 2.16.5.1 (3)-2,

it is highly probable that the wind velocity was almost zero.

3.4.2 Comparison of Flight profiles of the Aircraft between at the Time of the

Accident and based on flight manual

A comparison is made concerning the flight profiles of the Aircraft between at

the time of the accident as described in 2.1.1 and 2.16.5 and based on the flight manual

for the Aircraft as described in 2.14. The flight profile based on the flight manual is

computed by using the 0° flap takeoff procedures and the short field takeoff procedures

as described in 2.14.3 (4) and the takeoff roll distances and the takeoff distances for

each takeoff procedure as stated in 2.18.1. The computation is made on the assumption

that temperature is 34°C and takeoff weight is the maximum takeoff weight. The

comparison of height and speed for these factors is shown on Figure 3.4.2.

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Figure 3.4.2 Comparison of Flight Profile of the Aircraft between at the time

of the accident and based on flight manual

3.4.3 Takeoff Roll and Takeoff (10:57:13-10:57:41)

The takeoff point of the Aircraft at the time of the accident was approximately

630 m from the Runway 17 threshold as described in 2.1.1. On the other hand, on the

assumption that the Aircraft started the takeoff roll at about 10 m from the Runway

17 threshold, from the descriptions stated in 2.18.1, the takeoff point is approximately

537 m from the Runway 17 threshold at the time of the short field takeoff and the

maximum takeoff weight, and is approximately 690 m at the time of the 0° flap takeoff.

The flight manual described in 2.14.3 (4) specifies that the lift-off speed is 69 kt

for the short field takeoff procedure and is 78 kt for the 0° flap takeoff procedure. It is

highly probable that the takeoff speed*11 at the time of the accident was approximately

73 kt.

*11 “Take-off speed” means the speed at which all wheels lift.

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Figure 3.4.2 shows increases in speed when the Aircraft had the takeoff roll. It

shows that an increase in speed was lower than that for the short field takeoff where

flaps are set at 20°. It is somewhat likely that the reason for this is that the takeoff

weight of the Aircraft exceeded the maximum takeoff weight as described in 2.6.4; the

pitch angle fluctuated up and down at the center of the runway after starting the

takeoff roll as described in 2.1.1; and drag increased because the nose wheels lifted at

500 m from the Runway 17 threshold at a speed of approximately 65 kt.

As for the reason why the nose wheels lifted at a speed of approximately 65 kt,

judging from the facts that the speed at that time was lower than the lift-off speed

determined for the short field takeoff and the pitch angle fluctuated up and down at

the center of the runway, it is somewhat likely that the position of the C.G. was near

the aft limit as described in 3.3.4.

3.4.4 Speed at the Time of Takeoff

As described in 2.1.1, the Aircraft took off at 10:57:41 at the 630 m point of the

runway (remaining distance: 170 m). As described in 3.4.1 (1), judging from the

descriptions that the Aircraft stopped on the runway once and then started the takeoff

roll, the flaps at that time were set at 10° and the takeoff speed was 73 kt, it is probable

that the Captain performed the takeoff procedure at intermediate speed between the

lift-off speed of 78 kt for the 0º flap takeoff procedure and the lift-off speed of 69 kt for

the short field takeoff procedure as described in 2.14.3(4), performed the short field

takeoff procedure at 10º flap takeoff procedure; or he selected the 0º flap takeoff

procedure but reacted and took off because the Aircraft was approaching the runway

threshold.

3.4.5 State of Climbing just after Takeoff (10:57:42-10:57:54)

As shown in Figure 3.4.2, the climb angle just after takeoff during the flight of

the accident was almost the same as those for the short field takeoff and the 0° flap

takeoff.

As for the speed at that time, if the climb just after takeoff follows the short field

takeoff procedures, the Aircraft is expected to reach 80 kt before flying over 50 ft

obstacles, then retract gears and accelerate up to 90 kt while raising flaps. If the climb

follows the 0° flap takeoff procedures, the Aircraft is expected to reach 91 kt before

flying over 50 ft obstacles.

Figures 2.16.5.3-1 and 2.16.5.3-2 show that the rate of climb was about 500 fpm

from just after takeoff to climbing to 80 ft, meaning that the speed was reduced by 5-

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10 kt while climbing to 80 ft. The maximum speed after takeoff was approximately 76

kt. The speed was reduced to nearly 70 kt when reaching 80 ft.

It is somewhat likely that in the state that nose-up is easily to occur, because

the position of the C.G. was close to the aft limit, the Captain prioritized climbing than

speed, or the Captain was too late to recognize the decreased speed because, based on

his flight experience with the Aircraft, he thought that it was possible to accelerate

and climb even under such climbing situation. However, it was not possible to

determine why such flight was continued with the speed decreased because the

Captain died.

It is somewhat likely that under the situation just after takeoff, if the Captain

prioritized responding to the decreased speed and made the nose down, the speed

might not have reduced with decreasing the rate of climb and the flight could have

been continued. However, as shown in Figure 2.16.5.3-3, it is probable that the captain

continued to climb with excessive nose-up attitude, which decreased the speed, causing

the Aircraft to go into the backside flight described in 2.18.5 and reduced the speed at

which it was difficult to continue the flight.

3.4.6 Risks of Takeoff Slower than Lift-off Speed

The flaps for the Aircraft were set at 10° at the time of the accident. As

described in 2.18.1, there are no differences in takeoff performance between 0° and 10°

flaps. In this case, it is probable that the Aircraft must have accelerated up to the lift-

off speed determined in the 0° flap takeoff procedure before takeoff.

The lift-off speed for the 0° flap takeoff procedure is determined to be 78 kt.

However, the Aircraft took off at approximately 73 kt, a speed lower than that lift-off

speed. The Airplane Flying Handbook described in 2.18.4 states that an attempt to

take off at low speed and climb too steeply may cause aircraft to settle back to the

runway or make contact with obstacles. It is somewhat likely that during the takeoff

and climb, it took off slower than lift-off speed and climbed with excessive nose-up

attitude in such a way as to decrease the speed, and thereby the Captain could not

accelerate sufficiently to reach necessary climb speed. It is probable that these were

the factors for the subsequent decrease in height and the crash.

There is a high possibility that changing procedures prescribed in the flight

manual based on pilots’ own judgments affect operation safety. Such change should

never be made.

When it is difficult to accelerate up to the lift-off speed specified for selected

takeoff procedure, takeoff must be aborted without hesitation.

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3.4.7 Situation following climb immediately after Takeoff (10:57:55-10:58:00)

As described in 2.1.1, after reaching height approximately 90 ft, the Aircraft

was flying for about five seconds while descending gradually.

It is probable that continuing the climb with excessive nose-up attitude in such

a way as to decrease the speed, caused the Aircraft to get into the backside flight and

repeated up and down motion of the nose occurred. It is probable that the Aircraft

decreased the speed and was about to fell into power-on stall*12 condition.

3.4.8 Situation from Height Reduction to Crash (from 10:58:00)

As described in 2.1.1, the Aircraft rolled to the left at 10:58:00 and then

descended while gliding further to the left. It is probable that the speed just before the

Aircraft began to roll was approximately 62 kt. For estimated weight and 10° flaps,

the Stall Speed for this model is approximately 66 kt and the power-on stall(at 75 % of

maximum power) is calculated to be approximately 62 kt. Judging from these, it is

probable that the Aircraft could barely fly until around 10:58:00, but stalled and lost

height after that.

Further on that, it is probable that the Aircraft slightly regained speed as the

rate of descent increased and it came into shallow descent again just before crash.

As for the fact that the Aircraft veered to the left, it is probable that controlling

the Aircraft was difficult as a result that the Aircraft failed to climb and lost speed and

that it became unable to completely correct the characteristics of reciprocating single-

engine aircraft, which trend to veer to the left, with rudder or other operations.

Judging from the damage situations of the houses A, B, C and D and the

situation of the Aircraft after the crash described in 2.4.2, it is highly probable that,

just before the crash, the Aircraft made contact with the television antenna of house A

and then collided with the roof of house B. After that, the Aircraft crashed onto house

D, with its airframe turned upside down, without making contact with house C. From

this conditions, it is probable that the Aircraft took a nose-up attitude when it made

contact with the television antenna of house A and house B. Judging from the breakage

situation of the roof of house B and the damage situation of the body described in 2.11

(1), it is probable that the bottom side of the body crashed onto house B. It is probable

that the Aircraft bounded by the impact from the crash, took a forward turn movement

with a slight right twist during the bound, crossed over house C and crashed with the

airframe upside down.

*12 “Power-on stall” means a loss of speed at high power output.

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Figure 3.4.8 Estimated Aircraft Movement at the Time of Crash

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3.5 Analysis based on Mathematical Model

The flight status of the Aircraft at the time of the accident was reproduced by

calculations based on airplane characteristics data (mathematical model) analysis, and

was compared with the flight profile of the Aircraft at the time of the accident estimated

from images described in 2.1.1 and 2.16.5.

3.5.1 Conditions at the Time of the Accident

To analyze the flight status of the Aircraft at the time of the accident, when

conducting a simulation based on the developed mathematical model, the situation of

the Aircraft and the wind conditions at the time of the accident were assumed as below:

(1) Engine Power at Takeoff

As described in Figure 2.16.5.7-1, the manifold pressure was approximately 39

to 40 inHg from the takeoff roll to after takeoff. Therefore, it was assumed that

manifold pressure was 39 inHg as the value which has a big influence on the engine

power.

The temperature was assumed to be 34°C as described in 3.4.1 (4). In addition,

because the temperature on the runway was 38.1°C when the aviation meteorological

observation value was 34°C as described in 2.9.3, the temperature on the runway was

assumed to be 38°C at the time of the accident.

The engine power calculated from the performance chart described in 2.15.1 by

using these estimated values was 310 HP (89 % of 350 HP, the maximum rated power

of the same engine, hereinafter the same applies), and this value was assumed to be

the engine power of the Aircraft.

(2) Flap Position

Because it is highly probable that the flap position was 10° as described in 3.4.1

(2), the flap position was assumed to be 10°.

(3) Wind Conditions

It is highly probable that the wind at the time of the accident did not affect the

flight performance of the Aircraft and that there was almost no wind as described in

3.4.1 (4). Therefore, it was assumed that there was no wind.

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3.5.2 Mathematical Model of the Aircraft

The mathematical model of the Aircraft was developed to conduct analysis by

simulation. The flap position was set 10° as described in preceding paragraphs. The

aerodynamic coefficients with landing gears are retracted that was used for simulation

is shown in Figure 3.5.2. Parameters other than the aerodynamic coefficient indicated

in Figure 3.5.2 are shown in Table 3.5.2. The aerodynamic coefficient and parameters

shown in Figure 3.5.2 and Table 3.5.2 were calculated based on the information

provided by the manufacturer of the Aircraft and results of the flight test using the

type of aircraft.

The airplane

characteristics data was

developed by reflecting the

results obtained from the

flight test using the type of

aircraft based on the data

provided by the

manufacturer of the

Aircraft and is consistent

with the performance value

calculated by the

performance chart of the

flight manual under the same conditions. However, the obtained airplane

characteristics data does not completely match that of the Aircraft as that there can

be certain differences among individual aircraft.

Besides, the obtained results do not completely reproduce the flight status of

the Aircraft as the attitude and speed in the flight of the Aircraft estimated from

images described in 2.1.1 and 2.16.5 contain errors (mainly variation in values).

Figure 3.5.2 Relationship Diagram between Lift

Coefficient and Drag coefficient

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Table 3.5.2 Mathematical Model and Calculation Elements

Parameter Numerical Value

Coefficient of rolling friction 0.045

Increment of drag coefficient by

extending landing gears

0.023

Profile drag component of drag

coefficient

0.025

Ground effect*13

Induced drag: decreased by 50 % at 0 ft in

height decreased by 0 % at 25 ft in height

Aircraft mass 2,008 kg

Atmospheric density 1.14

Wing area 16.3 m2

3.5.3 Comparison at Takeoff Ground Rolls

The Aircraft stopped on the runway and started the Takeoff Ground Roll after

raising the engine RPM sufficiently as described in 3.4.1. As it is highly probable that

the flap position was set to be 10° as described in 3.5.1, it was compared with the

takeoff procedure at the 0° flap position during the takeoff ground roll. The test using

the type of aircraft verified that there was very little difference between 0° and 10°

flap positions in accelerating performance when applying the same procedure.

The heights and speeds of the Aircraft at the time of the accident and based on the

0° flap takeoff procedure by simulation are shown in Figure 3.5.3.

*13 “Ground effect” is the phenomenon that induced drag is reduced and the rate of the change of the

lift coefficient against the change of the attack angle increases as the situation of air flow around wings

change due to the influence of the ground when an aircraft is flying just close to the ground. Normally,

when the flight attitude is less than about the wing span, ground effect is thought to be observed. In

this case, as the lift drag ratio increases due to decreasing induced drag, the required thrust or the

required horsepower for the flight is small when the weight of the aircraft is the same.

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Figure 3.5.3 Comparison at the Takeoff Ground Rolls (0° Flap Takeoff

and the Takeoff of the Aircraft)

The flight of the Aircraft has noticeable differences as compared with the profile

of the 0° flap takeoff in the following two points:

・Slower acceleration at the takeoff ground roll (two-way arrow a in Figure 3.5.3)

・Earlier take-off (lower take-off speed) (two-way arrow b in Figure 3.5.3)

Based on the Takeoff Ground Roll distance of the 0° flap takeoff calculated from

the performance chart of the flight manual described in 2.18.1 and comparison

between the lift-off speed at the 0° flap takeoff of the flight manual and the flight

profile of the Aircraft at the time of the accident described in 2.14.3, the thrust and

engine power during the Takeoff Ground Roll were estimated.

The Takeoff Ground Roll Distance at the 0° flap takeoff is the distance from

where start of takeoff to where the main landing gear completely floats when the nose

of the Aircraft begins to rise at 78 kt; the speed at takeoff is 80.5 kt. Based on the

estimated flight profile at the time of the accident, assuming the Aircraft continued to

run to the point of 720 m where it is supposed to take off at the 0° flap takeoff, the

speed at the point was estimated to be 75.2 kt. The average acceleration was calculated

under the assumption that acceleration during the Takeoff Ground Roll was constant.

Then, the calculated average acceleration was converted to thrust and power. These

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values are shown in Figure 3.5.3.

Table 3.5.3 Estimation of Engine Power during the Takeoff Roll

State

Distance: 720 m

Speed at arrival (kt)

Average

acceleration

(m/s2)

Converted

to thrust

(%)

Converted

to power

(%)

0° flap takeoff 80.5 1.19 100 96

Takeoff of the Aircraft 75.2 1.04 87 78

It is found that the average acceleration at the takeoff of the Aircraft was 87 %

of the value at the 0° flap takeoff. Assuming this difference of the acceleration is caused

by decrease in thrust, the engine power was calculated to be 78 % (275 HP) based on

the performance chart of propellers.

3.5.4 Simulation at Takeoff and Climb

A simulation of the takeoff and climb of the Aircraft was conducted to match the

climb path by using the mathematical model of the Aircraft. In this simulation, the

engine power and flap setting of the Aircraft at the time of the accident assumed in

3.5.1 were used. In addition, the simulation was conducted under the assumption that

landing gear began to be retracted when the height exceeded 60 ft.

The simulation results when controlling an aircraft to match the climb path of

the Aircraft are shown in Figure 3.5.4.

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Figure 3.5.4 Comparison with the Simulation When Matching the Climb Path

As shown in these two figures, the flight of the Aircraft has noticeable

differences at the takeoff climb path as compared with that of the simulation in the

following two points:

・ Larger climb angle (two-way arrow c in Figure 3.5.3)

・ Decelerating during climb (two-way arrow d in Figures 3.5.3 and 3.5.4)

On the other hand, the results of the simulation show that an aircraft will not

decelerate and is supposed to accelerate to climb after the height exceeds

approximately 90 ft even if it retraces the same climb path of the Aircraft at the time

of the accident at the engine power assumed in 3.5.1 as shown in Figure 3.5.4.

The Aircraft took off at 10: 57: 41 and continued to climb for approximately 10

seconds. The climb rate during this time was 400 to 450 fpm and 3.0° to 3.5° when

converted to path angles. In addition, the speed decelerated by five to 10 kt in

approximately 10 seconds.

The climb performance of the Aircraft with the landing gears extended at 10°

flap and a speed of 70 kt was calculated from the aerodynamic coefficient and propeller

performance chart of the type of aircraft, and the resulting values are shown in Figure

3.5.4. No ground effects were taken into consideration in this calculation.

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Table 3.5.4: Estimation of Engine Power after Takeoff

Power (HP) Climbing performance (fpm) Converted to path angles (°)

335 590 4.6

310 520 4.0

280 420 3.3

220 190 1.4

The flight path angle of the Aircraft during climb is nearly the same as the climb

performance at 280 HP. However, it is probable that the average power during climb was

lower than that as the Aircraft was decelerating during this time.

3.5.5 Flight during Shallow Descent

The Aircraft was further decelerating as it was gradually descending from

10:57:55 to 10:58:00. Video images showed that the pitch angles repeatedly changed

as described in 2.1.1. In addition, it is highly probable that pitch angles repeatedly

changed within the range of a few degrees considering the relationship between the

height and speed at this time described in 2.16.5.3. The speed during this time was 65

to 70 kt, and it is probable that the Aircraft was flying backside described in 2.18.5.

The results of calculations of engine power which is required for the Aircraft to

perform horizontal flight with the landing gear up are shown in Figure 3.5.5. It is

somewhat likely that there are errors of about 0.02 in drag coefficient and about 20

HP in power as accurate data on drag was not obtained in particular, 65 kt and 70 kt

are near stalling speed.

Table 3.5.5 Engine Power Required for Horizontal Flight

Speed (kt) Power (HP)

65 190 (54 %)

70 160 (46 %)

75 135 (39 %)

80 125 (36 %)

It is somewhat likely that the power of the Aircraft which could not maintain

horizontal flight at the speed of 65 to 70 kt was 190 HP or less as described in Table

3.5.5. It is somewhat likely that large errors will occur in calculations that assume

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balanced steady state because the movement of the Aircraft was highly non-steady

especially during this phase.

The analysis based on the mathematical model in 3.5.3, 3.5.4, and this section

showed that the flight of the Aircraft (acceleration at takeoff, climb path, climb

performance, and speed during climb) could be reproduced, assuming there was a

decrease in engine power. Therefore, it is somewhat likely that the engine power of the

Aircraft at the time of the accident decreased.

3.6 Analysis on Engine Power of the Aircraft

As analysis using the mathematical model described in 3.5 showed the

probability of decreased engine output of the Aircraft, factors that are likely to give

influence on engine power were analyzed.

Factors that decrease engine power include flying operation, environmental

influence such as outside air temperature, and engine malfunction.

3.6.1 Decreased Engine Power due to Flying Operation

Influence of flying operation that leads to decreased engine power includes

throttle operation and the use of the air conditioner.

(1) Influence of Throttle Operation

As described in 2.16.5.7, the manifold pressure of the Aircraft during the takeoff

roll and takeoff climb was approximately 39 to 40 inHg according to the manifold

pressure gauge in the image taken within the Aircraft. As mentioned in 3.5.1 (1), it is

assumed that the engine power at this time was 310 HP (89 %). The engine power is

approximately 331 HP if manifold pressure is 42 inHg under the same condition.

Therefore, it is probable that engine power decreased by approximately six percent

because of low manifold pressure.

For throttle operation at the start of the takeoff roll, some pilots may set

manifold pressure slightly low in order to prevent overboost by the rise of manifold

pressure when the aircraft speed increases.

The manifold pressure of the Aircraft was approximately 39 to 40 inHg. It was

not possible to determine whether this manifold pressure of approximately 39 to 40

inHg was set by the throttle operation of the Captain or caused by other factors since

the captain was dead.

(2) Influence of the Use of Air Conditioner

According to the statements of the Passengers described in 2.1.2, it is highly

probable that the air conditioner was operated after the engine start. If the checklist

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of the flight manual was observed, the air conditioner might turned off before takeoff.

On the other hand there is some possibilities that the air conditioner was continuously

used during takeoff.

As described in Attachment 6, the teardown inspection of the air conditioner,

the use of an air conditioner decreases engine power by approximately one percent,

and the result of the teardown inspection showed that it was not possible to determine

the influence since the use of the air conditioner during takeoff was not able to be

identified based on the adhesion situation and the heat situation of the refrigerant oil

the Aircraft crashed upside down and was influenced by the subsequent fire.

3.6.2 Influence of Outside Air Temperature

According to the aeronautical weather observations described in 2.9.2, the

temperature at Chofu Airport around the time of the accident was 34°C. In addition,

as described in 2.9.3, it is probable that the outside air temperature over the runway

was approximately 38°C when observed temperature was 34°C. As described in 2.15.1,

when the outside air temperature is high, the performance of engine deteriorates.

Compared to the case of the standard atmosphere (15°C), it is highly probable that

engine power decreased by approximately 3.4 % and approximately 4 % when the

outside air temperature was 34°C and 38°C, respectively.

As previously mentioned, when the outside air temperature is high, engine

power decreases. The influence of outside air temperature has been reflected in the

performance table of the flight manual shown in Attachment 3-1 to 3-4; therefore, by

checking this, it is possible to confirm the influence of outside air temperature on flight

in advance. The analysis described in 3.5 also included the influence of outside air

temperature.

3.6.3 Possibility of Decreased Engine Power Due to Other Factors

Based on the failure mode (malfunction) resulting in engine malfunction, the

situation of the Aircraft at the accident, and the subsequent investigation result, the

possibility of the occurrences of engine power decrease was verified.

Table 3.6.3 shows failure modes resulting in engine malfunction in relation to

this accident, and instruments possibly indicating abnormal values when each failure

mode occurs or items for which abnormality will be found by the conducted tests and

others are marked with ○.

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Table3.6.3 Failure Mode Resulting in Engine Malfunction

Failure mode/instruments indicating

abnormal values

Ma

nif

old

pre

ssu

re

Fu

el

flow

En

gin

e R

PM

Sou

nd

Vib

rati

on

Exh

au

st g

as

colo

r

Tea

rdow

n i

nsp

ect

ion

Fu

el

syst

em

Obstruction (from pump to regulator) ○

Obstruction (from regulator to

distributor)

Obstruction (from distributor to

nozzle)

Obstruction (nozzle) ○

Failure of engine driven fuel pump ○

Contamination of water ○ ○ ○

Contamination of foreign matters

Improper fuel/air ratio ○ ○

Operation of the mixture lever by the

pilot

Injector maladjustment

Com

bu

stio

n Detonation ○ ○

Ind

uct

ion

syst

em

Turbocharger malfunction ○

Throttle operation by pilot ○ ○

Obstruction/break of induction pipe

(between valve and distributor)

Obstruction/break of induction pipe

(between distributor and cylinder)

Failure of induction valve

Cam failure

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Oil

pre

ssu

re

Decrease of oil pressure ○ E

xh

au

st s

yst

em

Break of exhaust pipe ○

Ign

itio

n s

yst

em

Plug non-ignition (wet fouling, wear) ○

Magneto malfunction

Oth

er Failure of piston rod and other

mechanical parts

(1) Malfunction of Fuel System

Figure 2.16.5.7 (1) shows that the value of the fuel flow indicator at the start of

the takeoff roll was normal, and as a result of the engine teardown inspection after the

accident described in 2.16.1, there was no obstruction in the fuel flow divider. Besides,

as described in 2.16.5.6, the results of sound analysis during the flight of the Aircraft

revealed that engine rotation was approximately 2,500 rpm, which was the maximum

rating. Based on these facts, it is probable that possibility of malfunction of the fuel

system was low.

(2) Malfunction of Combustion

As a result of engine teardown inspection after the accident described in 2.16.1,

the piston had no sign of early combustion, and evidence of normal combustion was

found on the piston head. Therefore, it is highly probable that detonation did not occur.

(3) Malfunction of Induction System

In the case of malfunction of the induction system, it is probable that the

manifold pressure gauge indicates abnormal values. However, based on Figure

2.16.5.7 (1), the value of the manifold pressure gauge during the takeoff roll, after

takeoff and during gear up indicated 39 to 40 (the maximum value is 42 in), and it is

probable that this value cannot be identified as the value in the case of clear

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malfunction of the induction system.

(4) Malfunction of Oil Pressure and Exhaust System

With regard to malfunction of oil pressure and the exhaust system, pre-flight

check (engine run-up) and engine teardown inspection found no sign of malfunction in

any failure mode. Therefore, it is probable that the possibility of the occurrence of these

malfunction is low.

(5) Malfunction of Ignition System

With regard to malfunction of the spark plug, checkup before flight and engine

teardown inspection found no abnormality, and therefore, it is probable that the

possibility of the occurrence of spark plug malfunction is low.

Besides, with regard to malfunction of the magneto, the results of engine

teardown inspection after the accident described in 2.16.1 and magneto teardown

inspection described in 2.16.3 revealed severe heat damage due to the fire after the

accident. Therefore, it was not possible to determine the possibility of occurrence of

malfunction.

(6) Other Malfunction

The result of teardown inspection did not reveal any mechanical malfunction

related to the power transmission system including the piston rods.

3.6.4 Verification Result of Engine Power

According to the analysis based on the mathematical model described in 3.5, it

is somewhat likely that the engine power of the Aircraft decreased.

The engine of the Aircraft was not in the state where evidence or signs showing

malfunction of the engine were able to be found due to damage from the impact of

crash and the subsequent fire, or there is a possibility that evidence and signs

themselves disappeared, however, from the analysis of investigation results related to

the engine described in the preceding paragraphs, it was not possible to obtain any

results clearly showing the occurrence of engine malfunction; hence, it was not possible

to determine that engine power decreased due to factors other than high outside air

temperature and low manifold pressure.

3.6.5 Relation between Decrease of Engine Power and RPM

As describe in 3.5, analysis based on the mathematical model showed a possible

decrease of engine power. If engine power decreases, as described in 2.7.2 (5), the

propeller pitch angle does not become shallower than the minimum value restricted

by low pitch stop despite the decrease of engine power, which leads to decrease in

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engine RPM itself . The decrease in engine RPM is indicated by the RPM indicator on

the instrument panel and is obvious by the change of engine sound, and therefore

pilots are able to easily recognize it.

Sound analysis described in 2.16.5.6 revealed that the engine RPM of the

Aircraft was maintained at approximately 2,460 rpm until the crash. It is probable

that there was no possibility of occurrence of decreased engine power under the

situation where engine RPM did not decrease. However, if the low pitch stop setting

was lower (shallower) than the specified value, it is somewhat likely that engine RPM

did not decrease despite the decrease of engine power.

As described in 2.15.2, Propeller Owner's Manual of propeller manufacturer

warns that adjustment of low pitch stop setting to lower value may hide any trouble

of engine power. If low pitch stop of the Aircraft was set at a value lower than the

specified value, it is somewhat likely that the Captain would not have been able to

recognize any decrease of engine power.

The teardown inspection of the engine, propeller and other parts described in

Attachment 6 revealed that there was a mark of the propeller pitch angle being 13° to

14° at the time of the crash on the preload plate. Besides, verification of high pitch

stop and low pitch stop after the teardown inspection described in Attachment 6

showed that the set value of low pitch stop was likely to be approximately 15°.

However, as shown in the propeller teardown inspection report described in

Attachment 6, in the record of overhaul immediately before the said propeller was

installed in the Aircraft (the year of 2005), low pitch stop was set at the specified value

of 17.6°, as described in 2.6.3 (3), there was no record of adjustment of low pitch stop

setting in the subsequent maintenance record, and propeller teardown inspection

reports state that “Although pre-load plate impact marks appeared to be below the low

pitch stop setting of 17.6°, they were likely caused by the impact forces and moments

forcibly twisting the blades beyond/lower than the low pitch stop setting.” Therefore,

it was not possible to determine whether low pitch stop setting was lower (shallower)

than the specified value.

3.6.6 Constant Decrease of Value Indicated on TIT Indicator

According to the records of flight tests in the past associated with periodical

maintenance work of the Aircraft (Attachment 5), the values indicated on the TIT

indicator had been lower than the normal operating range since 2012.

As described in 2.18. 8, the TCD and the revised TCD related to the malfunction

to degrade accuracy in values indicated on the TIT indicator was issued applicable to

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the TIT probe of the type of aircraft. Besides, the maintenance record of the Aircraft

included the record of maintenance work in accordance with the TCD and the revised

TCD but did not include the record of corrective work of low values indicated on the

TIT indicator at the flight test associated with periodical maintenance work. Therefore,

it is probable that the relevant corrective work was not implemented.

Consequently, it is somewhat likely that a full understanding of reasons for and

the background to the technical information of manufacturers and others was not

reflected in maintenance work and that flight test records after periodical

maintenance were not utilized in maintenance work to the full extent.

3.7 Preparation for Partial Power Loss of an Engine

As described in 2.18. 9, for partial power loss of an engine of single-engine

aircraft, in foreign countries, especially Australia and New Zealand, the information

related to preventive measures and operational reaction for the partial power loss of

an engine at takeoff and immediately after takeoff is collected and the results of

analyses of the collected information are summarized and published.

The preventive measures and operational reaction for the partial power loss of

an engine are summarized as follows;

(Preventive measures)

(1) Pre-flight planning

(2) Preflight check and inspection

(3) Preflight self-briefing

(4) Training

(Operational action for the partial power loss of an engine)

(1) Continue flight operation as far as possible for successful landing

(2) Decide about whether to return to the departure airport

It is probable that safety for flight of small private aircraft can be further

improved by positively preparing for an engine failure at takeoff and considering such

measures in advance.

3.8 Response of the Civil Aviation Bureau of the Ministry of Land,

Infrastructure, Transport and Tourism

The Civil Aviation Bureau of the Ministry of Land, Infrastructure, Transport and

Tourism as the safety actions to prevent reoccurrence of this accident , has implemented

concrete measures including rechecking of confirmation procedure regarding takeoff

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weight and the like before departure for operators of small aircrafts to enforce the

thoroughness to observe the laws and ordinances, procedures, has issued the Reminder

document to request reports the measures, and has conducted aviation safety seminars

where it has planned thorough confirmation before departure by pilot in command. In

addition, the Civil Aviation Bureau has strengthened the safety audit for air careers

based in Chofu Airport and the educational activities for operators of small aircrafts

through the meetings of operators of the Airport, and has enhanced the cooperative

relationship with the provider and administrator of airports at/from which small

aircrafts arrive/depart (including the provider and administrator of Chofu Airport (Tokyo

Metropolitan Government)) through exchange of opinions and information.

It is desirable for the Civil Aviation Bureau to enhance the relationship with the

provider and administrator of Chofu Airport (Tokyo Metropolitan Government) through

continued exchange of opinions and information, to comprehend progress made in

preventive measures by the administrator, and to give advice and guidance for the

administrator to ensure effort for preventive measures on a timely basis.

3.9 Response of the Provider and Administrator of Chofu Airport (Tokyo

Metropolitan Government)

In an explanation meeting with local residents held on June 2016, the provider

and administrator of Chofu Airport (Tokyo Metropolitan Government) explained the

safety actions taken after the accident, as follows; thorough confirmation before

departure by pilot in command; participation of maintenance engineers of Chofu Airport

in workshops for maintenance engineers organized by the national government;

mandatory participation of captains, maintenance engineers and aircraft dispatchers in

aviation safety seminars to further increase safety awareness; holding of regular

meetings for strengthening the system of responsibility during an emergency by

checking which organization is responsible at the time of and after an accident, along

with the provider and administrator of Chofu Airport (Tokyo Metropolitan Government)

plannings concerning a variety of support to provide rapid relief for injured parties, such

as actively working on people responsible for dealing with accident aircraft during an

emergency in liaison with relevant agencies, implementing a consultation service for

injured parties that will respond carefully to their requests and consultations, and

securing temporary homes for injured parties and taking responsibility for removal of

damaged houses. It is desirable for the provider and administrator (Tokyo Metropolitan

Government) to steadily implement these measures.

Moreover, the provider and administrator (Tokyo Metropolitan Government)

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appropriately monitors Chofu Airport and the surrounding area to prevent the

construction, planting and installation of tall buildings, plants and other structures that

affect arrival and departure of aircraft in the area. Furthermore, Chofu Airport has

received estimation results stating that, if the current runway end safety area is

maintained and used, any damage caused by any aircraft accident can be kept to a

minimum and there will be no cause for any accidents. It is probable that damage will

be limited to a certain level, even if undershoot and overrun occur at Chofu Airport.

Therefore, it is desirable for the provider and administrator of Chofu Airport (Tokyo

Metropolitan Government) to notify the operators of information on the runway end

safety area at Chofu Airport through AIP or other methods as quickly as possible.

3.10 Improvement of Safety

In this accident, the small private Aircraft crashed into a residential area and

caused injury to residents as well as damages to houses around Chofu Airport. In order

to prevent from recurrence of this kind of accidents, measures for ensuring and

improving operation safety of small private aircraft were considered, as below.

3.10.1 Ensuring Safety of Small Private Aircraft

It is somewhat likely that the Captain of the Aircraft took off without following

the procedure prescribed in the flight manual described in 3.4.6 in excess of maximum

takeoff weight and without compliance with requirements for performance prescribed

in the flight manual. Compliance with the weight, speed and procedures prescribed in

the flight manual is important for operation safety. If the weight and the position of

the C.G. are not within the allowable range, the controllability, stability or flight

performance of aircraft will be reduced, and flight safety cannot be ensured. In that

case, flight must not be undertaken.

The Civil Aviation Bureau of MLIT needs to address the following items to

ensure the safety of small private aircraft.

(1) Confirmation before Departure by Pilot in Command

As described in 2.6.4 and 2.18.1, the Aircraft did not satisfy the requirements

for performance prescribed in the flight manual. that is, takeoff distance shall be

runway length in use or less, along with exceeding the maximum takeoff weight at the

time of the takeoff.

As described in 2.18.2, it is necessary for the captain at the confirmation before

departure provided in Article 164-14 of Ordinance of the Civil Aeronautics Act and

Article 73-3 of the Civil Aeronautics Act to confirm a fact that not only weight of an

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aircraft is within maximum takeoff weight but also it satisfies the requirements for

performance prescribed in the flight manual, however, it is probable that the

understanding about the importance is not sufficiently instilled among captains of

small private aircraft, differing from operators of air transport services who provide

the Operation Manual following the Detailed Guideline of Examination of Operation

Manuals..

Over the past five years, there have been two fatal accidents involving small

private aircrafts affected by inappropriate weight and the position of the C.G. of the

aircraft. In this accident, a small private aircraft caused an accident as it took off

without satisfying the requirements for performance prescribed in the flight manual

as well as exceeding the maximum takeoff weight.

Therefore, it is necessary to promote pilots of small private aircrafts to

understand the importance to confirm that requirements for performance prescribed

in the flight manual are satisfied, in addition to the importance to comply with

maximum takeoff weight and limit for the position of the C.G., as confirmation before

departure, at the occasions like specific pilot competency assessments and aviation

safety seminars.

Moreover, to prevent confirmation errors made by the captain, and ensure safety

flight, it is preferable to request persons with knowledge on aviation to double check

the records confirmed by the captain as much as possible.

(2) Assumption of Decreased Flight Performance at Takeoff

Along with compliance with the speed and procedure prescribed in the flight

manual, as for the actions to the situation of degraded flight performance leading to

the situations where the flight could not be continued due to lack of acceleration or

decrease in speed during takeoff, it is necessary to enforce instructions and trainings

to pilots of small private aircraft to plan the actions in advance including to follow the

emergency procedure prescribed in the flight manual and confirm these actions

thorough self-briefing by a pilot himself at the time of preparation before departure.

3.10.2 Improvement of Safety at Airports

As described in 3.4.4, it is somewhat likely that the Aircraft took off as it was

approaching the runway threshold.

As described in 2.18.6 (4), in Japan, there are cases where entrance taxiways

are connected to extensions of runways as seen in Osaka International Airport and

others, and enable a maximum use of runway length. In this manner, regarding

making maximum use of runway length at takeoff, it will allow a pilot to have a margin

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to make a decision during takeoff roll and contribute to improving safety, it is

necessary to study and compile the cases of effective measures connecting entrance

taxiways to runway thresholds in order to make maximum use of runway length and

inform aerodrome providers and administrators of these case studies.

3.10.3 Measures for Increasing Safety in Small Private Aircraft Maintenance

As described in 2.18.6, the TCD and the revised TCD had been issued to the

Aircraft to correct TIT probe failure, which decreased the accuracy of values indicated

on the TIT indicator, and the maintenance records of the Aircraft showed that the

required measures had been taken. However, this investigation found that the reading

of the TIT indicator had been constantly low in the past flight test records.

For this reason, proper maintenance measures should have been applied to deal

with the decreased values indicated on the TIT indicator, but it is probable that the

measures were not taken. Therefore, it is somewhat likely that a full understanding

of the reason for and the background to the technical information of the manufacturers

and others was not reflected in maintenance work and that the flight test records after

periodical maintenance were not utilized in maintenance work to the full extent.

It is necessary for small private aircraft to be securely maintained based on a

proper understanding of technical information. Therefore, appropriate management of

technical information and implementation of maintenance work based on the

information is important to ensure aircraft safety.

4. CONCLUSION

4.1 Findings

(1) Takeoff Weight and Balance

When the accident occurred, it is highly probable that the weight of the Aircraft

is estimated to have been approximately 2,008 kg and that the Aircraft was

approximately 58 kg heavier than the maximum takeoff weight, which was 1,950 kg.

Moreover, it is probable that the position of the C.G. was +146.0 to +146.5 in. aft of the

reference line, and it was close to the aft limit at the time of the maximum takeoff

weight. (3.3)*14

It is probable that the weight and the position of the C.G. of the Aircraft were

*14 The number at the end of each paragraph in this section (4. Conclusion) refers to a main subsection

number of “3 Analysis” where a relevant description can be found.

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not confirmed sufficiently by the Captain before departure. (3.3.1)

It is not possible to determine whether the Captain recognized the fact that the

weight of the Aircraft exceeded the maximum takeoff weight before departure, it is

somewhat likely that he had insufficient understanding of the risks of making flights

under such situation and insufficient safety awareness of observing laws, regulations

and provisions. (3.3.2)

If an aircraft exceeding its maximum takeoff weight engages in flight, the

aircraft has reduced takeoff and climb performance. Besides, when the position of the

C.G. is close to the aft limit, the condition will result in excessive nose-up attitude and

decreased controllability, stability or flight performance during a low-speed flight,

which may lead to an unexpected stall. (3.3.3)

In this accident, it is also highly probable that the excess of the Aircraft weight

deteriorated the takeoff and climb performance, and in the state that the nose-up was

easily to occur because the position of the C.G. close to the aft limit, was leading to the

excessive nose-up attitude at the time of takeoff and climb, and decreased

controllability, stability and flight performance during a low-speed flight, which

brought the Aircraft to a condition easy to stall. It is highly probable that these were

the factors that led the Aircraft to take off at low speed, have the excessive nose-up

attitude, and enter stall.

When planning flights, captains must accurately identify the weight of persons,

baggage on board and fuel, calculate the weight and the position of the C.G. based on

these factors by using calculating documents or other materials, and be sure to confirm

that the weight and the position of the C.G. are within allowable ranges. (3.3.4)

The amount of fuel to be loaded should be carefully examined when there is a

possibility that the weight of the aircraft is likely to exceed the maximum takeoff

weight or the weight restricted from the airplane performance. (3.3.4)

(2) Flight of the Aircraft at the Time of the Accident

It is probable that the Aircraft underwent standing takeoff, and it is highly

probable that the flap setting were 10°. It is highly probable that the temperature was

34°C and there was almost no wind. (3.4.1)

POH/AFM specifies that the lift-off speed is 78 kt for the 0° flap takeoff

procedure, but it is highly probable that the takeoff speed at the time of the accident

was approximately 73 kt.

As for the fact of the Aircraft showed the movement of lifting the nose gear wheel

at 500 m point from the threshold of the runway 17 with speed of approximately 65 kt,

it is somewhat likely that because the position of the C.G. was close to the aft limit.

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(3.4.3)

With regard to takeoff at the speed of approximately 73 kt, it is somewhat likely

that the Captain performed the takeoff procedure at intermediate speed between the

lift-off speed of 78 kt for the 0º flap takeoff procedure and the lift-off speed of 69 kt for

the short field takeoff procedure, performed the short field takeoff procedure with the

10º flap setting, or he selected the 0º flap takeoff procedure but reacted and took off

because the Aircraft was approaching the runway threshold. (3.4.4)

As for the reasons why a climbing was continued with the speed decreased, it is

somewhat likely that in the state that nose-up is easily to occur, because the position

of the C.G. was close to the aft limit, the Captain prioritized climbing than speed, or

the Captain was too late to recognize the decreased speed because, based on his flight

experience with the Aircraft, he might have thought that it was possible to accelerate

and climb even under such climbing situation. It is probable that the captain continued

to climb with excessive nose-up attitude, which decreased the speed, causing the

Aircraft to go into the backside flight and reduced the speed at which it was difficult

to continue the flight. (3.4.5)

It is somewhat likely that the Aircraft took off slower than lift-off speed and

climbed with excessive the nose-up attitude, which decreased the speed, and thereby

the Captain could not accelerate sufficiently to reach necessary climb speed. It is

probable that these were the factors for the subsequent decrease in height and the

crash.

When it is difficult to accelerate up to the lift-off speed specified for selected

takeoff procedures, takeoff must be aborted without hesitation. (3.4.6)

As for the situation following climb immediately after takeoff, it is probable that

continuing the climb with excessive nose-up attitude in such a way as to decrease the

speed, caused the Aircraft to get into the backside flight, and led it to be about to fell

into power-on stall condition. (3.4.7)

It is probable that the Aircraft could barely fly until around 10:58:00, but stalled

and lost height after that. It is probable that the Aircraft took a nose-up attitude when

it crashed onto a house, bounded by the impact from the crash, took forward turn

movement with a right twist during the bound, and crashed with the airframe upside

down. (3.4.8)

(3) Analysis based on Mathematical Model

Simulations were conducted by taking account of temperature and manifold

pressure at the time of the accident and by calculating engine power from the manual

of the engine manufacturer. The results of the simulation showed that an aircraft will

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not decelerate and is supposed to accelerate to climb after the height exceeds

approximately 90 ft even if it retraces the same climbing route of the Aircraft at the

time of the accident. (3.5.4)

Assuming there was a decrease in engine power before the Aircraft started

shallow descent, acceleration at takeoff, climb path, climb performance, and speed

during climb of the Aircraft was able to be reproduced. Therefore, it is somewhat likely

that the engine power of the Aircraft at the time of the accident decreased. (3.5.5)

(4) Analysis on Engine Power of the Aircraft

Factors that decrease engine power include flying operation, environmental

influence such as outside air temperature, and engine malfunction. (3.6)

From the analysis of investigation results related to the engine, it was not

possible to obtain any results clearly showing the occurrence of engine malfunction,

and it was not possible to determine that engine power decreased due to factors other

than high outside air temperature and low manifold pressure. (3.6.4)

(5) Response of the Civil Aviation Bureau of the Ministry of Land, Infrastructure,

Transport and Tourism

It is preferable for the Civil Aviation Bureau of the Ministry of Land,

Infrastructure, Transport and Tourism to enhance the relationship with the provider

and administrator of Chofu Airport (Tokyo Metropolitan Government) through

continued exchange of opinions and information, to comprehend progress made in

preventive measures by the administrator, and to give advice and guidance for the

administrator to ensure effort for preventive measures on a timely basis. (3.8)

(6) Response of the Provider/Manager of Chofu Airport (Tokyo Metropolitan

Government)

It is desirable for the provider and administrator of Chofu Airport (Tokyo

Metropolitan Government) to steadily implement measures explained at the meeting

with local residents on June 2016, and notify the operators of information on the

runway end safety area at Chofu Airport through AIP or other methods as quickly as

possible. (3.9)

(7) Improvement of Safety

It is necessary for the Civil Aviation Bureau of MLIT to promote pilots of small

private aircraft to understand the importance to confirm that requirements for

performance prescribed in the flight manual are satisfied, in addition to the

importance to comply with maximum takeoff weight and limit for the position of the

C.G., as confirmation before departure, at the occasions like specific pilot competency

assessments and aviation safety seminars.

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Furthermore, along with compliance with the speed and procedure prescribed

in the flight manual, as for the actions to the situation of degraded flight performance

leading to the situations where the flight could not be continued due to lack of

acceleration or decrease in speed during takeoff, it is necessary to enforce instructions

and trainings to pilots of small private aircraft to plan the actions in advance including

to follow the emergency procedure prescribed in the flight manual and confirm these

actions thorough self-briefing by a pilot himself at the time of preparation before

departure. (3.10.1)

Regarding making maximum use of runway length at takeoff, it will allow a pilot

to have a margin to make a decision during takeoff roll and contribute to improving

safety, it is necessary to study and compile the cases of effective measures connecting

entrance taxiways to runway thresholds in order to make maximum use of runway

length and inform aerodrome providers and administrators of these case studies.

(3.10.2)

4.2 Probable Causes

It is highly probable that this accident occurred as the speed of the Aircraft

decreased during takeoff and climb, which led the Aircraft to stall and crashed into a

residential area near Chofu Airport.

It is highly probable that decreased speed was caused by the weight of the Aircraft

exceeding the maximum takeoff weight, takeoff at low speed, and continued excessive

nose-up attitude.

As for the fact that the Captain made the flight with the weight of the Aircraft

exceeding the maximum takeoff weight, it is not possible to determine whether or not

the Captain was aware of the weight of the Aircraft exceeded the maximum takeoff

weight prior to the flight of the accident because the Captain is dead. However, it is

somewhat likely that the Captain had insufficient understanding of the risks of making

flights under such situation and safety awareness of observing relevant laws and

regulations.

It is somewhat likely that taking off at low speed occurred because the Captain

decided to take a procedure to take off at such a speed; or because the Captain reacted

and took off due to the approach of the Aircraft to the runway threshold.

It is somewhat likely that excessive nose-up attitude was continued in the state

that nose-up tended to occur because the position of the C.G. of the Aircraft was close to

the aft limit, or the Captain maintained the nose-up attitude as he prioritized climbing

over speed.

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Adding to these factors, exceeding maximum takeoff weight, takeoff at low speed

and continued excessive nose-up attitude, as the result of analysis using mathematical

models, it is somewhat likely that the decreased speed was caused by the decreased

engine power of the Aircraft; however, as there was no evidence of showing the engine

malfunction, it was not possible to determine this.

5. SAFETY ACTIONS

5.1 Safety Actions Taken after the Accident

5.1.1 Safety Actions Taken by the Civil Aviation Bureau of the Ministry of Land,

Infrastructure, Transport and Tourism

(1) Issuing the reminder document to small aircrafts operators

Immediately after the Accident, concrete measures including rechecking of

confirmation procedures regarding takeoff weight and the like in order to enforce the

thoroughness to observe the laws and the ordinance procedures were implemented,

and the reminder document to request reports regarding the measures were issued.

(2) Thorough Implementation of Confirmation by Captains before Departure

It was decided to put more effort into training sessions to encourage captains to

obtain weather information, prepare flight plans (include the confirmation of weight

and balance of aircraft), obtain basic knowledge about engine test run and conduct

accurate confirmation.

a Lecturers are dispatched to training sessions for pilots, such as aviation

safety seminars, to enhance the quality of these events. In addition, the

staff of the Civil Aviation Bureau holds lectures to educate captains about

the importance of thorough confirmation and preflight safety inspection.

Adding to this, classes are held regarding the necessities to confirm takeoff

weight, C.G. position and balance based on the flight manual at these

seminars.

b Leaflets (Include necessities to confirm takeoff weight, C.G. position and

balance prior to a departure) are distributed to explain how to ensure the

safety of small aircrafts to pilots who will take specific pilot competency

assessment given by pilot competency examiners, in order to educate them

about flight safety.

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(3) Promoting to Conduct Accurate Maintenance by Private Aircraft Maintenance

Engineer

At training sessions, the staff of the Civil Aviation Bureau informs and educates

private aircraft maintenance engineer about the importance of complying with aircraft

manuals, relevant laws and regulations, and of conducting proper maintenance.

(4) Purchase of Aviation Insurance Policy for Private Aircraft

Leaflets are distributed to pilots who will take specific pilot competency

assessment given by pilot competence examiners, in order to encourage them to

purchase an insurance policy. Adding to this, when private aircrafts use a temporary

helipad and national airports, the coverage status of aircraft/pilots is confirmed prior

to a flight under conditions of taking-out an aviation insurance (third party liability

insurance) in order not to fly without appropriate insurances. And instruct to take the

same measures at other airports than the ones managed by a national government.

(5) Provision of Information Services

In order to increase the use of existing flight information services for pilots in

flight, materials describing the outline of “TCA advisory service” and “area/en-route

information service (AEIA)” were prepared and distributed at training sessions for

small aircrafts pilots.

(6) Appropriate Acquisition of Business License

Instructions to acquire a business license were reinforced, and the provision of

educational information and information about license acquisition for businesses

involving aircraft was enhanced.

a Educational leaflets about the acquisition of a business license were

provided to private aircraft pilots and aircraft owners via regional Civil

Aviation Bureaus and industry associations.

b Measures were taken to make it easier to acquire a business license using

flow charts and application forms necessary in business license acquisition

procedures.

(7) Deployment of Measures for Business Operators who Use Small Aircraft

Air carrier services and business operators who use small aircraft were also

informed of measures (1) and (2) for private aircraft, even though they have clearly

stipulated their procedures for confirmation before departure and aircraft

maintenance in their operation and maintenance manuals.

(8) Reinforcement of Cooperation with Providers and administrators of Airports

where Small Aircrafts Takeoff and Land

(i) Regular Information Exchange with Airport Managers

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a It was decided that continued discussion with persons in charge of

administration of the Chofu Airport associated with the provider and

administrator of Chofu Airport (Tokyo Metropolitan Government) will be

made.

b It was decided that discussion with persons in charge of administration

of all airport will be held at training sessions provided by the Aerodrome

Support and Aeronautical Service every year.

(ii) Training Sessions for Airport Administrators

It was decided that training sessions will be provided (in FY2015) to persons

in charge of airports administrated by local government or private organization

(including public heliports) to deepen their knowledge about SMS (safety

management system) in airports and improve airport safety.

(9) Organizing Operators of Small Aircrafts

It was decided to promote the organization of small aircrafts operators for each

airport in order to facilitate communication and thorough implementation of safety

measures.

a Recommendation was made to implement safety information sharing with

small aircrafts operators through the Airport Committee at the FY2015 airport

manager training session.

b Through relevant organizations, small aircrafts operators permanently

stationed at airports and private aircraft groups were invited to positively

participate in the Airport Committee meeting hosted by providers and

administrators of airports.

(10) Promotion of measures to take new safety actions and initiatives to raise safety

awareness

Civil Aviation Bureau of MLIT regularly holds “Safety Promotion Board relating

to small aircrafts and the like” since 2016 and while taking account of opinions

expressed by experts and relevant organizations at the board as carrying out

investigations and studies relating to establishments of Safety Actions for small

aircrafts, and promotes the initiatives to take new safety actions and to raise safety

awareness.

5.1.2 Safety Actions Taken by the Provider and administrator of Chofu Airport (Tokyo

Metropolitan Government)

(1) Aiming at the dissemination of and education about safety measures, on

January 20, 2016, a safety training session was held for business operators and private

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aircraft operators based there (19 companies and 2 individuals). Lecturers were

invited from the Civil Aviation Bureau of the Ministry of Land, Infrastructure,

Transport and Tourism, air carrier services and the organization of private aircraft

operators who were based at Chofu Airport. Participation in the session was on a

voluntary basis.

(2) Safety inspection of aircraft was carried out for all 10 business operators who

use aircraft for their business and are based at Chofu Airport. Inspection was

conducted three times (from July (after the accident) to September 2015, from October

to December 2015, and from January to March 2016).

5.2 Safety Actions Required by the Civil Aviation Bureau of the Ministry

of Land, Infrastructure, Transport and Tourism

(1) In this accident, small private aircraft crashed into a residential area and caused

injury to residents as well as damages to houses, however it is highly probable that

the aircraft was flying with exceeding the maximum takeoff weight and without

satisfying the requirements for performance prescribed in the flight manual. Therefore,

it is necessary to promote pilots of small private aircraft to understand the importance

to confirm that requirements for performance prescribed in the flight manual are

satisfied, in addition to the importance to comply with maximum takeoff weight and

limit for the position of the C.G., as confirmation before departure based on the

provision of Article 164-14 of Ordinance of the Civil Aeronautics Act, at the occasions

like specific pilot competency assessments and aviation safety seminars.

Furthermore, along with compliance with the speed and procedure prescribed

in the flight manual, as for the actions to the situation of degraded flight performance

leading to the situations where the flight could not be continued due to lack of

acceleration or decrease in speed during takeoff, it is necessary to enforce instructions

and trainings to pilots of small private aircraft to plan the actions in advance including

to follow the emergency procedure prescribed in the flight manual and confirm these

actions thorough self-briefing by a pilot himself at the time of preparation before

departure.

(2) As stated in 3.4.4, it is somewhat likely that the Aircraft took off as it was

approaching the runway threshold. Regarding making maximum use of runway length

at takeoff, it will allow a pilot to have a margin to make a decision during takeoff roll

and contribute to improving safety, study and compile the cases of effective measures

connecting entrance taxiways to runway thresholds in order to make maximum use of

runway length and inform aerodrome providers and administrators of these case studies.

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6. RECOMMENDATIONS

It is highly probable that this accident occurred as the speed of the Aircraft

decreased during takeoff and climb, which led the Aircraft to stall and crashed into a

residential area near Chofu Airport.

It is highly probable that decreased speed was caused by the weight of the Aircraft

exceeding the maximum takeoff weight, takeoff at low speed, and continued excessive

nose-up attitude.

As for the fact that the Captain made the flight with the weight of the Aircraft

exceeding the maximum takeoff weight, it is not possible to determine whether or not

the Captain was aware of the weight of the Aircraft exceeded the maximum takeoff

weight prior to the flight of the accident because the Captain is dead. However, it is

somewhat likely that the Captain had insufficient understanding of the risks of making

flights under such situation and safety awareness of observing relevant laws and

regulations.

It is somewhat likely that taking off at low speed occurred because the Captain

decided to take a procedure to take off at such a speed; or because the Captain reacted

and took off due to the approach of the Aircraft to the runway threshold.

It is somewhat likely that excessive nose-up attitude was continued in the state

that nose-up tended to occur because the position of the center of gravity of the Aircraft

was close to the aft limit, or the Captain maintained the nose-up attitude as he

prioritized climbing over speed.

Adding to these factors, exceeding maximum takeoff weight, takeoff at low speed

and continued excessive nose-up attitude, as the result of analysis using mathematical

models, it is somewhat likely that the decreased speed was caused by the decreased

engine power of the Aircraft; however, as there was no evidence of showing the engine

malfunction, it was not possible to determine this.

In this accident, small private aircraft crashed into a residential area and caused

injury to residents as well as damages to houses, however the Aircraft was flying with

exceeding the maximum takeoff weight and without satisfying the requirements for

performance prescribed in the flight manual, and over the past five years, there have

been two fatal accidents involving small private aircraft affected by inappropriate weight

and position of the center of gravity of the aircraft ( (1) Mooney M20C, JA3788, which

crashed when landing at Yao Airport in March 2016, and (2) Cessna 172N Ram, JA3814,

which veered off the runway of Otone Airfield, Kawachi Town, Inashiki-gun, Ibaraki

Prefecture, and made a fatal contact with a ground worker in August 2012). In view of

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the result of these accident investigations, as operation safety of small private aircraft

needs to be improved, the Japan Transport Safety Board recommends the Minister of

Land, Infrastructure Transport and Tourism pursuant to Article 26 of the Act for

Establishment of the Japan Transport Safety Board to take the following measures:

(1) Promote pilots of small private aircraft to understand the importance to confirm

that requirements for performance prescribed in the flight manual are satisfied, in

addition to the importance to comply with maximum takeoff weight and limit for the

position of the center of gravity, as confirmation before departure, at the occasions like

specific pilot competency assessments and aviation safety seminars.

Enforce instructions and trainings to pilots of small private aircraft to plan the

actions in advance including to follow the emergency procedure prescribed in the flight

manual and confirm these actions thorough self-briefing by a pilot himself at the time

of preparation before departure. along with compliance with the speed and procedure

prescribed in the flight manual, as for the actions to the situation of degraded flight

performance due to lack of acceleration or decrease in speed during takeoff.

(2) Study and compile the cases of effective measures connecting entrance taxiways

to runway thresholds in order to make maximum use of runway length and inform

aerodrome providers and administrators of these case studies as maximum use of

runway length at takeoff, will allow a pilot to have a margin to make a decision during

takeoff roll and contribute to improving safety.

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Figure 1 Three View Drawings of Piper PA-46-350P

3.44

13.11

8.81

Unit: m

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- 124 -

Figure 2 Seating Chart

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Attachment 1-1 Photos taken from inside

of the Aircraft (Right Wing and Flaps) (i)

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Attachment 1-2 Photos taken from inside

of the Aircraft (Right Wing and Flaps) (ii)

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Attachment 2 Inside of the Aircraft

(instrument panel)

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Attachment 3-1 Performance Chart (Takeoff

Ground Roll Distance — 0° Flaps)

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Attachment 3-2 Performance Chart (Takeoff

Distance — 0° Flaps)

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Attachment 3-3 Performance Chart (Short Field

Takeoff — Takeoff Ground Roll Distance)

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Attachment 3-4 Performance Chart (Short

Field Takeoff — Takeoff Distance)

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Attachment 4 Wind data Around the Time of

the Accident

Time of observation Instantaneous wind

direction (true bearing)

Instantaneous wind speed

(kt)

10:57:12 243 1

10:57:15 236 1

10:57:18 211 1

10:57:21 200 0

10:57:24 175 0

10:57:27 184 0

10:57:30 184 0

10:57:33 164 0

10:57:36 152 0

10:57:39 172 1

10:57:42 184 1

10:57:45 178 1

10:57:48 199 1

10:57:51 216 1

10:57:54 218 1

10:57:57 204 1

10:58:00 181 1

10:58:03 191 1

10:58:06 179 1

10:58:09 139 1

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Attachment 5 Flight Test Record (Extract)

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Attachment 6 Teardown Inspections of Engine,

Propeller and Others

1 Teardown Inspection of Engine and Propeller

In order to investigate conditions of the engine and propeller of the Aircraft at the

time of crash, teardown inspections were conducted at the engine manufacturer’s facility

(in the U.S.) from January 12 to 13, 2016. The inspections of the engine and engine

accessories was performed by the engine manufacturer and that of the propeller,

propeller governor and turbochargers was performed by the propeller manufacturer.

1.1 The main contents described in the engine manufacturer’s report obtained from the

U.S. investigative agency are as follows:

(Excerpt)

(1) Engine as Received or First Viewed

The engine had thermal signatures consistent with a post crash fire. The engine did

not have intercoolers and the propeller was attached. The engine mount and nose gear

structure were also still attached to the engine.

(2) Engine Data

Last Annual Inspection by: unknown Date: 4/27/15

Last Overhaul by: Lycoming Engines Date: 3/22/03

Compression Test: N/A

Comments: could not rotate engine due to thermal damage and rust in cylinder.

Valve Action:Rotate Engine if possible to verify continuity through engine.

Comments: could not rotate engine due to thermal damage and rust in cylinder.

① Basic Information

The engine history shows the engine left the Lycoming Factory on March 22,

2003 and was shipped to Van Bortel Aircraft, Inc. The engine was returned to

Lycoming March 24, 2005 for a prop strike inspection at 73.5 hours. The engine was

disassembled on March 29, 2005 for the inspection and was reassembled, tested and

returned to the customer.

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② Propeller and Propeller Governor

As first viewed by this ASI, the propeller remained attached to the engine via

the crankshaft flange. The propeller was removed from the engine and all

disassembly of the propeller was completed by the Hartzell propeller ASI. The

propeller governor was found secure to its mount on the crankcase and was removed.

The propeller governor gasket screen was clear of contamination or debris. The

propeller governor was disassembled by the Hartzell propeller ASI.

③ Fuel System

As first viewed by this ASI, the fuel system was thermally breached consistent

with a post crash fire. All fuel lines showed evidence of thermal damage. The fuel

servo remained secure to the oil sump and air inlet housing assembly. The throttle

cable was cut prior to arrival at Lycoming.

The fuel flow divider, hard fuel lines, and fuel injector nozzles were found secure

and were removed. The fuel flow divider hard lines were flow checked for

obstructions using low pressure air with no obstructions noted. The single piece fuel

nozzles were removed from the individual cylinders and visually checked for

obstructions. Debris consistent with combustion materials and rust were found in

the nozzles consistent with debris created during the post crash fire and subsequent

firefighting efforts. The flow divider was removed from the case and disassembled.

The diaphragm had signs of thermal stress but was not breached. The mechanic

breached the diaphragm in an attempt to remove the diaphragm from the flow

divider body. No fuel was found in the fuel flow divider. The engine driven geared

fuel pump was found secure to its mount and was removed. Signs of thermal damage

consistent with a post crash fire were observed on the outer case of the fuel pump

and the inlet and outlet fuel lines were thermally breached. The engine driven

geared fuel pump was rotated by hand. Suction and compression could not be

confirmed due to the pump hanging up during rotation consistent with thermal

damage from the post crash fire.

④ Magneto

The magnetos were found secure on their respective mounts with heavy thermal

damage noted to the rear of each magneto. As such, the magnetos were removed but

confirmation of spark could not be made due the thermal damage received on the

magnetos.

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⑤ Spark Plugs

As first viewed by this ASI, all the spark plugs remained secure in their

respective cylinder locations. All the spark plugs were removed and photographed

in their as-removed condition. The spark plugs were then cleaned and visually

inspected for center electrode wear and cracks in the center tower. No visual signs

of defect were noted.

⑥ Ignition Harness

The ignition harness was thermally destroyed consistent with damage received

during the post crash fire.

⑦ Starter

As first viewed the starter was found secure on its mount with Bendix gear

engaged on the starter ring gear. The starter was removed but was not tested.

⑧ Generator

As first viewed, both the left and right alternators were found secure on their

associated mounting surfaces and brackets. The left alternator destroyed by fire.

⑨ Vacuum Pump

Both vacuum pumps were destroyed due to thermal damage, no data tag

information was available.

⑩ Lubricating System

The oil suction screen was removed from the oil sump and visually inspected for

debris or metal with none found. The oil sump was removed and small amount of oil

was noted in the sump. It was a gray/black in color and was free of debris or metal.

The remote mount spin on oil filter was shipped with the engine. It showed thermal

damage consistent with a post crash fire and was not opened up.

⑪ Supercharging System

As first viewed, the turbochargers remained secure to their mounts on the

engine. The left turbocharger was thermally destroyed with the compressor housing

and compressor impellers thermally destroyed with only the shaft remaining. The

right turbocharger was able to rotate freely and had no damage to the compressor

impellers. The variable absolute controller was thermally damaged at the outer

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housing consistent with damage received during a post crash fire. There was no

Manifold Pressure Relief Valve found; however, most of the air inlet housing that

holds the Manifold Pressure Relief Valve was thermally destroyed consistent with

damage received during a post crash fire.

⑫ Push Rods, Accessory Housing and Oil Pump

Push Rod Tubes: #2 intake impact damage, all others normal

Push Rods: #2 intake impact damage, all others normal

Accessory Housing: Thermal signatures consistent with a post crash fire

Oil Pump: No contamination found inside the oil pump

Gear, Splines and Drivers: rust and thermal signatures were found on most of the

accessory housing gears, splines and shafts.

As first viewed by this ASI, the push rod tubes and push rods were attached to

the engine with thermal signatures consistent with a post impact fire. The #2 intake

push rod tube was impact damaged as well as the #2 intake push rod. With the

cylinders removed, each push rod was visually inspected and found to be

unremarkable. The rockers had normal wear patterns. The accessory housing was

removed and the oil pump disassembled. The gears in the accessory housing were

rust covered but the gears turned freely. The oil pump gears showed no evidence of

metallic particles or scoring on the oil pump housing.

⑬ Cylinders

As first viewed by this ASI, the cylinders remained secure on their respective

mounts on the crankcase. Thermal signatures consistent with a post crash fire were

evident on all six cylinders. Molten metal was found lodge in the cylinder cooling

fins on cylinder #1 and #3. All the cylinders were removed and piston removed from

the connecting rods. All cylinders exhibited rust inside on the cylinder walls with

cylinder #6 exhibiting the heaviest rust. The pistons exhibited no signs of detonation

and had normal combustion signatures on the piston head. The intake and exhaust

valves were not removed from the cylinders. Visual inspection of the valves heads

were unremarkable.

⑭ Crankcase

The crankcase was disassembled to remove the crankshaft and camshaft. All

bearing journals were normal and unremarkable. The propeller governor gear set

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screw was visually verified to be in place. The silk thread was visually verified to be

in place.

⑮ Crankshaft

As first viewed by this ASI, the crankshaft remained inside the crankcase and

would not rotate. As the cylinders were removed, the rust present on the crankshaft

bearings appeared to contribute to the resistance in rotation of the crankshaft. With

all the pistons removed, the crankshaft was rotated at the propeller flange and was

able to complete 360 degrees of rotation. All counterweights had the rollers installed

with the internal retaining ring in place. The counterweights and the rollers

exhibited heavy surface rust. The rear crankshaft gear bolt was installed with the

locking tab in place. The crankshaft flange run out was not checked.

⑯ Camshaft, Connecting Rods and Bearings

As first viewed by this ASI, the camshaft, connecting rods, and bearings

remained in place inside the crankcase. The crankcase was split to remove the

camshaft and the connecting rods were removed from the crankshaft. The camshaft

lobes were normal and did not show signs of pitting. The camshaft gear was normal

with rust present on the face of the gear. The bearings on the #6 connecting rod

showed heavy rust that did not allow the connecting rod to move freely and was

consistent with rust build up after a post crash fire. All tappet faces were normal

and showed no signs of pitting.

(3) Engine Disassembly Observations

On January 12th, 2016, the engine was moved to the NTSB/FAA teardown room

within the Lycoming Factory for disassembly. The engine was uncrated and placed on a

ring stand for initial observation and photo documentation. The engine was completely

disassembled and condition documented. The propeller, propeller governor, and

turbochargers were disassembled by an ASI from Hartzell Propeller.

The engine showed signs of a post crash fire that burnt over the entire engine and

propeller. There were no intercooler present and the engine mount and nose gear

structure remained attached at the engine mounts.

There was nothing found during the course of this engine tear down that would have

precluded this engine from making power prior to impact. All damage noted in the

sections above is consistent with the engine being involved in the post crash fire of the

aircraft at the accident site.

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1.2 Teardown Inspections of Propeller/Turbocharger

The main contents described in the propeller manufacturer’s report obtained from

the U.S. investigative agency are as follows:

(1) SUMMARY AND ANALYSIS OF FINDINGS

The propeller and propeller governor were examined at Lycoming Engines in

Williamsport, PA on January 12, 2016.

The propeller remained attached to the engine during the impact sequence and

subsequent recover and shipping. The propeller assembly, including spinner was

removed from the engine by Lycoming personnel before the teardown examination was

conducted. Impact marks on both blade preload plates indicated the propeller was on or

near the low pitch setting during the impact sequence. One blade exhibited chord wise

abrasion in the tip area on the camber side, indicating rotation before impact. The other

blade was mostly consumed by the post-crash fire.

A governor teardown examination did not reveal any anomalous conditions.

(2) CONCLUSIONS

The propeller showed signs of rotation with power ON and blade angles on or near

the low pitch stop during the impact sequence.

The pre-load plate markings indicated the propeller was operating on or near the

low pitch stop at impact. Blade damage (camber side scoring, aft bending and twisting

leading edge down), plus pitch change knob damage all indicate the blades impacted

while rotating and at a negative angle of attack, resulting in high twisting moments

towards low pitch. Although pre-load plate impact marks appeared to be below the low

pitch stop setting of 17.6°, they were likely caused by the impact forces and moments

forcibly twisting the blades beyond/lower than the low pitch stop setting. At low aircraft

speeds (<70 KIAS), negative blade angles of attack are experienced at power levels below

approximately 50% power.

There were no discrepancies noted that would prevent propeller operation prior to

impact. The propeller was over-serviced with grease which may have caused slow

response to power changes, or RPM oscillations, but would not have prevented operation.

All damage was consistent with high impact forces twisting the blades towards low pitch,

and heat damage due to the post-crash fire.

There were no discrepancies noted with the propeller governor that would prevent

normal operation prior to impact. The control arm/shaft assembly was missing and

appeared to have been forcibly removed either during the impact sequence or during

recovery/transporting. The governor otherwise appeared undamaged and could be

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rotated by hand.

(3) Propeller Teardown Report

① General Comments

This type propeller is a 2-blade single-acting, hydraulically operated, constant

speed model. Oil pressure from the propeller governor is used to move the blades to

the high pitch (blade angle) direction. A spring and blade twisting moment move the

blades toward the low pitch direction in the absence of governor oil pressure. The

blades and hub are of aluminum construction. Propeller rotation is clockwise as

viewed from the rear.

② Installation Data

Reference Hartzell Installation Data Sheet #874

(Data reference the 30-inch station)

Low Pitch: 17.6 + 0.2 degrees

High Pitch: 40.5 + 0.5 degrees

Note: Manufacturing records indicated this propeller model was originally

assembled with a low pitch angle of 16.0 degrees. However, the last

overhaul records in 2005 confirm the low pitch was set to17.6°before it was

reinstalled on the accident aircraft.

③ Service History:

According to JTSB investigators the propeller was replaced in 2005 after a

propeller object/ground strike and at the time of the accident had 1541.01 hours

since new.

S/N Date of manufacture TTSN TSO

Hub HA6 6/6/1988 1541.01 926.54

Blades H00909 6/13/1988 1541.01 926.54

F96066 6/13/1988 1541.01 926.54

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④ Propeller Serial Number HA6

⑤ Factory Serial Number: A12067

⑥ Blade Model: F8074K

Blade #1: S/N H00909

Blade #1: S/N F96066

⑦ Blade Orientation:

The blades were arbitrarily number 1-2 clockwise as viewed from the rear of the

propeller. The hub serial number was between the 1 and 2 blades.

⑧ As Received Condition:

The propeller was attached to the engine when shipped to Lycoming. The

propeller had been removed from the engine by Lycoming personnel prior to my

arrival. Photos #1 and #2 show the condition of the propeller prior to the

examination. The majority of one blade had been consumed by the post-crash fire.

The other blade showed bending in the aft direction and twisting leading edge down

(toward low pitch). The spinner dome and bulkhead remained attached to the

propeller but approximately 30% was consumed by the post-crash fire.

⑨ Spinner Dome

The spinner dome was still mounted to the spinner bulkhead but crushed on

one side and approximately 30% of it was consumed by the post-crash fire. The

forward bulkhead was present but had debonded and the flange was bent in several

Photo#1: Propeller (before teardown) Photo#2: Spinner part

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places.

⑩ Spinner Bulkhead:

The spinner bulkhead remained attached to the hub. Approximately 40% of the

bulkhead was either consumed or heat damaged by the post- crash fire.

⑪ Propeller Cycling

Propeller cycling was not attempted before disassembly due to lack of mounting

provisions in the facility, lack of blade paddles to rotate the blades, and/or a method

to pressurize the piston.

Neither blade could be rotated by hand force only.

⑫ Engine/Propeller Mounting

The propeller remained attached to the engine throughout the crash sequence

and subsequent recovery and shipping to Lycoming. The propeller was removed from

the engine at Lycoming.

The mounting flange and all six mounting studs appeared intact and

undamaged. There was no visually remarkable bending of the mounting studs or

damage to the counter bores. The O-ring appeared undamaged and there was oil

present in the hub bore leading up to the pitch change rod.

⑬ Cylinder (P/N: B-2428-1 Rev. AO, S/N: GK6881)

The cylinder was intact but showed signs of heat distress on the exterior surface

(charring, bubbling, etc.). The low pitch stop screw had been forced forward,

stripping mounting threads in the cylinder but remained in the cylinder and

required further unthreading to remove (see Photo #3).

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⑭ Piston:

The piston was intact and

unremarkable. The piston O-ring

was intact but showed signs of

heat distress (hard/not pliable,

flat and discolored in some

areas.)

There was oil and what

appeared to be sludge aft of the

piston which was later

determined to be blade bearing

grease. Grease and grease

byproducts from over-servicing

the blade bearings had migrated forward through the hub bushing to the area

behind the piston (see Photo #4).

Photo #5 shows the excessive amount of grease in the hub.

⑮ Pitch Change Rod

The pitch change rod was intact and unremarkable.

⑯ Fork

The fork was intact. The only visible damage was the plastic bumpers had

melted and charred causing some surface discoloration.

⑰ Spring

The spring was intact. The only remarkable trait was some oil coking/charring

Photo#3: Low Pitch Stop

Photo#4: Piston part Photo#5: Hub part

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discoloration from heat exposure during the post-crash fire.

⑱ Pitch Stops

Low Pitch Stop: The low pitch stop screw had been forced forward, partially out

of the cylinder, stripping mounting threads in the cylinder. But

it remained in the cylinder and required further unthreading to

remove (see Photo #3).

High Pitch Stop: The high pitch stop was intact and unremarkable.

⑲ Hub Assembly

The preload plate shelves showed deformation on the aft/trailing edge

quadrants consistent with the preload plate lip deformation. This deformation

indicates the impact forces were a combination of aft and opposite direction of

propeller rotation (see Photo #6). The hub was also excessively over-serviced with

bearing grease (see Photo #5).

⑳ Preload Plates: (see Photos #7 and #8)

NOTE: For this propeller model, when the blade knob is aligned with the hub

parting line, the blade angle at the reference station is 48°(knobφ

=12°+ 36 = 48° ). On a two-blade Compact series propeller, pitch

change knob impressions are sometimes made on the opposite-side

preload plate during the impact sequence. This impression mark can be

measured to provide some blade angle information.

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The # 1 preload plate had a knob impression mark at approximately 34°

corresponding to a blade angle of 14r (3.6esponding to a blade angle of 14°(below

the low pitch stop angle).

The # 2 preload plate had a knob impression mark at approximately 35°

corresponding to a blade angle of 13°(4.6°below the low pith stop angle).

㉑ Blade Bearings and Blade Pitch Change Knobs:

The blade bearings were intact but the plastic ball spacer and grease on both

blades had been heat compromised, was stiff and either crumbled or congealed

together. The plastic bushings on the pitch change knobs had melted. The pitch

change knob for blade #2 was fractured in the direction opposite low pitch with no

signs of fatigue in the fracture surface. The pitch change knob for blade #1 was bent

in the direction opposite low pitch with a visual crack at its base (Photo #9). There

were small ball indentations in the bearing races on the camber (forward) side of

each blade indicating impact forces in the aft direction.

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㉒ Propeller Blades: (see Photos #10 through #12)

There were impressions on each blade butt from pre-load plate contact

indicating forces in a helical path (i.e. propeller rotation + forward speed) (Photo

#10).

Blade #1

paint, camber side - Chordwise scoring, heat damage (scorched/

charred/corroded).

paint, flat side - Heat damaged (scorched/charred/corroded).

bend - Bent aft starting approximately 15roded).

charred/corrod

twist - Leading edge down/toward low pitch.

lead edge damage - Gouging, nicks, dents and deformation from tip

extending approximately 8”toward shank. De-ice

boot missing, assumed consumed by fire.

trail edge damage - Outboard 6°melted/consumed

knob condition - Visually bent opposite low pitch, cracked at base, pre-

load plate impact mark at shoulder.

Blade #2 (See Appendix A for material composition analysis of melted

material shown in Photo #12)

paint, camber side - Outboard ~24” of blade melted/consumed, shank

was scorched/charred/corroded

- Outboard 24”of blade melted/consumed, shank

was scorched/charred/corroded

bend - Not distinguishable.

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twist - Not distinguishable.

lead edge damage - Heat damaged, de-ice boot missing, assumed

consumed by fire.

trail edge damage - Heat damaged.

knob condition - Fractured, no visual signs of fatigue, only overload.

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(4) Propeller Governor Teardown Report

① Propeller Governor Model: V-11-1

② Governor Serial Number: G1455J

③ General Comments:

The propeller governor is an engine/propeller RPM sensing device and high

pressure oil pump. In a constant speed propeller system, the governor responds to a

change in engine/propeller RPM by directing oil under pressure to the propeller

hydraulic cylinder (when over speeding), or by releasing oil from the hydraulic

cylinder to the drain (when under speeding). The change in oil volume in the

hydraulic cylinder changes the propeller blade angle and returns the

engine/propeller RPM to the pilot selected value. The V-11 series governor is a

pressure-to-increase pitch, non-feathering, intermediate capacity governor. The V-

11-1 governor is specifically configured for use on the Piper PA-46-350P with the

TIO-540-AE2A engine.

④ Installation Specifications:

Refer to Hartzell Manual 130B for V-11-1 specifications

⑤ Service History:

No service history data for the governor was presented prior to the examination

⑥ As Received Condition:

The propeller governor was still mounted to the engine but the control arm/shaft

assembly was fractured and missing (see Photo #13). The high RPM stop/adjust

screw was intact. The governor was removed from the engine without difficulty and

the drive shaft turned freely by hand. The gasket screen did not have any

remarkable debris.

⑦ Governor Head and Spool Valve:

The governor head was removed without difficulty. It appeared the control

arm/shaft assembly was forcibly pulled from the head but the rack and rack adjust

screw remained in the head. The spool valve appeared undamaged and could be

moved by hand within the governor body. Internal oil discoloration implied evidence

of high temperatures inside the governor head.

⑧ Flyweight Unit:

The flyweights were intact and moved freely on their pivot points. The flyweight

disk turned freely with shaft rotation by hand.

⑨ Oil Pump and Governor Body:

The oil pump gears were undamaged and turned freely by hand. The oil pump

cavity was free of any foreign objects or debris. The pressure relief valve was intact.

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See Photo #14 for the governor components disassembled and examined.

(5) Metallurgical Laboratory Report

Photo – Melted Propeller Blade

① Test Method: The blade was sectioned and chemical analysis was conducted.

② Test Summary: The melted blade was sectioned and appeared porous throughout.

Chemical evaluation indicates that the material is an aluminum alloy containing

copper, magnesium, and silicon, but due to the burned nature of the blade, some

the constituents could have burned off or changed in amount.

③ Test results:

Chemical analysis of the melted propeller indicate the blade was produced from

an aluminum alloy containing copper, magnesium, and silicon.

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Table Chemical Analysis Result

Element Melted blade 2025 AA Limi

Copper 3.3 3.9 to 5.0

Magnesium 1.11 0.050 max.

Silicon 1.13 0.50 to 1.20

Iron 0.28 1.00 max.

Manganese 0.47 0.40 to 1.20

Nickel 0.01 ---

Zinc 0.27 0.25 max.

Titanium 0.04 0.15 max.

Chromium 0.02 0.10 max.

Other, ea. <0.05 0.05 max.

Other, total <0.15 0.15 max.

Aluminum Rem. Remainder

(6)Turbocharger Teardown Report

① TURBO SYSTEM GENERAL INFORMATION:

The turbocharger system installed on the TIO-540-AE2A engine used on the PA-

46-350P aircraft is called a “Variable Absolute Pressure System.” The variable

absolute pressure controller senses deck pressure, compares it to a reference

absolute pressure, and adjusts the wastegate butterfly via the turbo controller valve

(controlling turbocharger speed) to maintain the desired horsepower at varying

altitudes. It differs from the non-variable version, however, in that the turbo

controller is directly linked to the engine throttle, and through a system of cams and

followers, adjusts itself to varying power settings, achieving the optimum deck

pressure for a given throttle movement. A pressure relief valve set slightly in excess

of maximum deck pressure is provided to prevent damaging overboost in the event

of a system malfunction. A sonic venturi (customer supplied) is incorporated to

provide a constant source of compressed air to the cabin pressurization system. An

intercooler (customer supplied) is added to cool the compressor outflow and increase

cylinder charge air density. The system includes the components listed in Table 1.

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Only the two turbochargers and wastegate valve body were recovered from the

accident aircraft and presented for examination. The controller and pressure relief

valve were either not recovered or consumed by the post-crash fire and therefore not

examined.

② Wastegate:

The wastegate valve was free and had full range of motion. The wastegate

actuator body was completely consumed by the post-crash fire; only the valve

housing assembly, actuator shaft and return spring remained. See Photo #1.

Photo #1 – Wastegate valve

③ Right Turbo (Turbo #1 ) Information

There was light impact damage, some evidence of heat/sooting, dirt and oily

surface on turbine side. There was no evidence of eroding, fretting, or damage to

attachment/mounting surfaces.

Center Hub Rotation Assembly (CHRA):

Note: Dimensional inspections of the CHRA assembly components were not

performed since the turbocharger easily rotated freely by finger force;

teardown was deemed unnecessary.

The compressor wheel spun with the turbine wheel on the shaft. It had

moderate blade damage.

The turbine wheel appeared undamaged. BackPlate: Intact

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Photo #2 – Remains of compressor wheel Photo #3 – Turbine wheel

④ Left Turbo (Turbo #2) Information

The left turbo exhibited extensive fire damage and corrosion. There was no

evidence of eroding, fretting, or damage to attachment/ mounting surfaces. The

compressor housing, compressor wheel and mounting surfaces were consumed

by fire.

Clearance Between T-Wheel blades and Housing: Not at all blade locations.

Clamps and Lock plates: All turbine lock plates and fasteners intact. The

compressor clamps and fasteners were missing or

consumed by the post- crash fire.

Condition of Compressor Housing: Consumed by post-crash fire.

Center Hub Rotation Assembly (CHRA):

NOTE: Dimensional inspections of the CHRA assembly components were

not performed either due to extreme heat damage or since

teardown was conducted at Lycoming; the proper tooling and

fixtures were not available.

Center Bearing Housing:

・Existence of Residual Oil. Dry and extreme corrosion present.

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Photo #4 – As-Received Condition 1

⑤ Conclusions and Additional Comments:

The Right turbocharger assembly was consistent with the design data and

there was no evidence of mechanical malfunction. The turbocharger appeared

functional and the wheels spun freely with light finger force. There was

evidence of lubrication present (wet surfaces and free rotation).

The Left turbocharger assembly components that were not consumed by

the post-crash fire appeared consistent with the design data and there was no

evidence of mechanical malfunction prior to impact or the post-crash fire. No

liquid indication of lubrication was present, presumably due to post-crash fire

heat and time elapsed from the accident. The compressor wheel and compressor

housing were almost completely consumed by the post-crash fire indicating the

level of heat exposure.

Although both turbochargers did not show any evidence of pre-impact

mechanical malfunction, the turbocharging system performance cannot be

verified due to the missing turbo controller, wastegate actuator and pressure

relief valve. None of the turbocharger components presented showed evidence

Photo #5 Remains of compressor wheel Photo #6 – Turbine

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of wheel/housing rub but based on the lack of external impact damage, does not

indicate the turbochargers were not turning at impact. There is no evidence to

suggest the turbocharger system was not functional prior to impact.

2 Teardown Inspection of Magneto

In order to investigate the ignition system of the engine, a teardown inspection of

the magneto, which is the main equipment of the system, was performed on August 2,

2016. Since this equipment was charred— especially the non-metal part was

significantly consumed and damaged—it was not possible to perform visual inspection,

functional inspection and other inspections. See Photos #21 and #22.

2.1 General Comments

Starting with a teardown of one magneto and visually checking the inside, it was

significantly consumed by the post-crash fire and it was not possible to perform a

functional test. Then the other magneto was tore down and the condition was similar.

The main conditions confirmed at the time of the teardown are as follows:

(1) The ignition harness was mostly consumed;

(2) The distributor block was damaged by fire;

(3) The distributor rotor was melted;

(4) The coil assembly was damaged by fire;

(5) The field through capacitor was consumed;

(6) Although the breaker point was damaged by fire, there was no major damage

at the contact point which could be inspected;

(7) The primary lead wire connecting the field through capacitor and the coil

assembly was not broken.

(8) The rotating magnet could be rotated freely by hand.

Photo #21: Magneto (Right side) Photo #22: Magneto (Left side)

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3 Teardown inspection of Air Conditioning System (Air Conditioner)

According to the flight manual of the Aircraft, it is prohibited to use the air

conditioning system (hereinafter referred to as “Air Conditioner”) at the time of takeoff

and landing because it cannot obtain normal takeoff climb performance if it is on. The

aircraft manufacturer said that the engine power is reduced by approximately 1 % (3-5

HP) when the Air Conditioner is operating. In addition, the use of the Air Conditioner

when an aircraft leaves a parking area will not cause any safety problem if it is turned

off before takeoff. The compressor of the Air Conditioner recovered from the accident site

was tore down and the state of the Air Conditioner before the accident was investigated.

To investigate the operating condition of the Air Conditioner at the time of takeoff,

a teardown inspection was performed by the Air Conditioner manufacturer on May 9,

2016. The main details of the report prepared by the manufacturer are as follows:

3.1 Teardown of Each Part of the Compressor

(1) Electromagnetic Clutch

Attraction (electric power ON) and separation (electric power OFF) operations of the

clutch plate were available and there was no problem.

(2) Around the Shaft Rotor

A. Front thrust bearing

No wear or deformation was found and there was no problem.

B. Bearing rolling part of main bearing/shaft roller

No wear or deformation was found and there was no problem.

C. A set of rear thrust bearing

No wear or deformation was found and there was no problem.

(3) Around the Planet Plate

A. A set of ball bearings/a set of gears

No wear or deformation was found and there was no problem.

B. Rod sliding part of the planet plate

There was no problem.

(4) Piston / Cylinder Bore Part / Rod Sliding Part of the Piston

No wear or deformation was found and there was no problem.

(5) Around Valves

A. Discharge valve

No gap was found and there was no problem.

B. Inlet valve

No gap was found and there was no problem.

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3.2 Investigation on Whether Air Conditioner was Operating before the Accident

(1) Result of analysis of black substance attached inside the crank chamber

A black substance was found attached to the lower inner surface of the crank

chamber after the accident. No substance was found on the upper side.

Fourier transform infrared light absorption spectrometry was performed to

identify the black substance attached and it matched refrigerating machine oil, SP10A.

The refrigerating machine oil was found attached only on the lower side of the wall

surface in the crank chamber after the accident.

Photo #23: a-a Cross Section

事故後下側

正規搭載位置での下側This is the lower side of the part if the Aircraft was not

upside-down

This was the lower side after the accident

(1) Electromagnetic Clutch

(2) Shaft Rotor

(3) Planet Plate

(4) Piston

(5) Discharge

Valve/Inlet Valve

Figure. Attachment 6 3.2 (2) Compressor (cross-section)

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(2) Observation

The table below shows the condition of refrigerating machine oil attached to the

inner wall of the crank chamber in relation to the operating condition of the air

conditioning system from departure to takeoff and at the time of takeoff.

Table. Attachment 6 3.2 (2) Observation Related to the Operating Condition of Air

Conditioning System

Operating condition of

air conditioning

system

Condition of refrigerating machine oil attached to the inner surface of the

crank chamber

Befo

re

tak

eoff

At

the

tim

e o

f

tak

eoff

At the time of takeoff At the time of the

accident

After the fire

(heating)

1

Op

era

tin

g

Op

era

tin

g

Attached to the entire

inner surface

(Oil stirred by movable

parts)

Attached to the entire

inner surface

(Attached to the inner

surface of the crank

chamber due to oil’s

viscosity)

Burned refrigerating

machine oil (black

substance) attached

to the entire inner

surface 2

Op

era

tin

g

Not

Op

era

tin

g

Attached to the entire

inner surface

(Attached due to oil’s

viscosity)

3

Not

Op

era

tin

g

Op

era

tin

g

Attached to the entire

inner surface

(Oil stirred by movable

parts)

4

Not

Op

era

tin

g

Not

Op

era

tin

g

Attached to the entire

inner surface

(Oil and condensed

refrigerant were mixed

and the viscosity was

reduced)

Attached to the entire

inner surface

(The part turned upside

down after the accident,

resulting in the

attachment of oil and

condensed refrigerant to

the lower surface due to

gravity)

It is somewhat likely

that burned

refrigerating machine

oil (black substance)

attached to the lower

inner surface after

the accident

4 Inspection of High and Low Pitch Stops of Propeller

After the teardown inspection by the propeller manufacturer, angles of high and low

pitch stops were inspected from the propeller hub of the Aircraft with the cooperation of

relevant administrative organization in Japan from June 10, 2016. Main details of the

report are as follows:

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Although angles of high and low pitch stops were inspected from the recovered

propeller hub of the Aircraft, it is possible that propeller pitch angles measured this time

do not completely represent the condition before the accident because the hub was

deformed due to the crash and heat.

Under normal conditions, the blade angle of the propeller matches the value

measured at the 75 % position of the radius. However, both propeller blades of the

Aircraft were damaged and it was difficult to measure them. Then, by measuring the

operating range of the Pitch Change Nob (hereinafter referred to as “Nob”) attached to

the butt of the propeller blades to change the propeller pitch angle, values of high and

low pitch stops were calculated. According to the material from the propeller

manufacturer, the following relationship exists between angle of Nob φ and propeller

pitch angle β:

β= 48-φ

Based on the technical data and other materials provided by the propeller

manufacturer, the butt part of the propeller blade, including Nob, was created by a 3D

printer (see Photo #24). In addition, since the adjusting bolt of the low pitch stop was

damaged, it was substituted by a bolt with the same length (see Photo #25).

The conditions of assembled parts measured are shown in Photos #26 to #29, and

the results of Nob angle measurements are shown in the table below.

Photo #24: Butt part of propeller

blade created by 3D printer

Photo #25: Low Pitch Stop Bolt (bottom)

and Substitute Bolt (top)

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Table. Attachment 6 4 Propeller Pitch Angles

Blade #1 (deg) Blade #2 (deg)

Nob φ Pitch β Nob φ Pitch β

High Pitch Stop 7.5 40.5 7.7 40.3

Low Pitch Stop 33.0 15.0 32.8 15.2

Photo #27 Low Pitch Stop Blade #1 Photo #26: High Pitch Stop Blade #1

Photo #28: High Pitch Stop Blade #2 Photo #29: Low Pitch Stop Blade #2